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LC REGULATION: BEFORE SERVICING 
OR REPRODUCING ANY PART OF THIS 
DOCUMENT, ALL CLASSIFIC ATION 
MARKINGS MUST BE CANCELLED. 


SUMMARY TECHNICAL REPORT 


OF THE 



NATIONAL DEFENSE RESEARCH 


This document contains information affecting the national defense of the 
United States within the meaning of the Espionage Act, 50 U. S. C., 
31 and 32, as amended. Its transmission or the revelation of its contents 
in any manner to an unauthorized person is prohibited by law. 

This volume is classified CONFIDENTIAL in accordance with security 
regulations of the War and Navy Departments because certain chapters 
contain material which was CONFIDENTIAL at the date of printing. 
Other chapters may have had a lower classification or none. The reader 
is advised to consult the War and Navy agencies listed on the reverse of 
this page for the current classification of any material. 


Manuscript and illustrations for this volume were pre- 
pared for publication by the Summary Reports Group of the 
Columbia University Division of War Research under con- 
tract OEMsr-1131 with the Office of Scientific Research and 
Development. This volume was printed and bound by the 
Columbia University Press. 

Distribution of the Summary Technical Report of NDRC 
has been made by the War and Navy Departments. In- 
quiries concerning the availability and distribution of the 
Summary Technical Report volumes and microfilmed and 
other reference material should be addressed to the War 
Department Library, Room 1A-522, The Pentagon, Wash- 
ington 25, D. C., or to the Office of Naval Research, Navy 
Department, Attention: Reports and Documents Section, 
Washington 25, D. C. 


Copy No. 

239 


This volume, like the seventy others of the Summary Tech- 
nical Report of NDRC, has been written, edited, and printed 
under great pressure. Inevitably there are errors which 
have slipped past Division readers and proofreaders. There 
may be errors of fact not known at time of printing. The 
author has not been able to follow through his writing to 
the final page proof. 

Please report errors to: 

JOINT RESEARCH AND DEVELOPMENT BOARD 
PROGRAMS DIVISION (STR ERRATA) 

WASHINGTON 25, D. C. 

A master errata sheet will be compiled from these reports 
and sent to recipients of the volume. Your help will make 
this book more useful to other readers and will be of great 
value in preparing any revisions. 



SUMMARY TECHNICAL REPORT OF DIVISION 3, NDfcC^&fe,. 


VOLUME 1 






* 




«%> 




ROCKET AND UNDERWATER 


ORDNANCE 


LCJREGJtJLATION^ BEFORE SERVICING 
OR REPRODUCING ANY PART OF THIS 

document, all^classification 
markings must BE~CANCElEETT 


office of scientific research and development 

YANNEVAR BUSH, DIRECTOR 

national defense research committee 

JAMES B. CONANT, CHAIRMAN 

DIVISION 3 
F. L. HOVDE, CHIEF 


WASHINGTON, D. C., 1946 


• f* f * ... 

NATIONAL DEFENSE RESEARCH COMMITTEE 

■ ■ * i • • • • i 

James B. Conant, Chairman 
Richard C. Tolman, Vice Chairman ■' : • 

Roger Adams Army Representative 1 

Frank B. Jewett Navy Representative 2 

Karl T. Compton Commissioner of Patents 1 

Irvin Stewart, Executive Secretary 


1 Army representatives in order of service : 

Maj. Gen. G. V. Strong Col. L. A. Denson 

Maj. Gen. R. C. Moore Col. P. R. Faymonville 

Maj. Gen. C. C. Williams Brig. Gen. E. A. Regnier 

Brig. Gen. W. A. Wood, Jr. Col. M. M. Irvine 

Col. E. A. Routheau 


2 Navy representatives in order of service : 

Rear Adm. H. G. Bowen Rear Adm. J. A. Furer 
Capt. Lybrand P. Smith Rear Adm. A. H. Van Keuren 
Commodore H. A. Schade 

3 Commissioners of Patents in order of service: 
Conway P. Coe Casper W. Ooms 


NOTES ON THE ORGANIZATION OF NDRC 


The duties of the National Defense Research Committee 
were (1) to recommend to the Director of OSRD suitable 
projects and research programs on the instrumentalities 
of warfare, together with contract facilities for carrying 
out these projects and programs, and (2) to administer 
the technical and scientific work of the contracts. More 
specifically, NDRC functioned by initiating research 
projects on requests from the Army or the Navy, or on 
requests from an allied government transmitted through 
the Liaison Office of OSRD, or on its own considered ini- 
tiative as a result of the experience of its members. Pro- 
posals prepared by the Division, Panel, or Committee for 
research contracts for performance of the work involved 
in such projects were first reviewed by NDRC, and if 
approved, recommended to the Director of OSRD. Upon 
approval of a proposal by the Director, a contract per- 
mitting maximum flexibility of scientific effort was ar- 
ranged. The business aspects of the contract, including 
such matters as materials, clearances, vouchers, patents, 
priorities, legal matters, and administration of patent 
matters were handled by the Executive Secretary of 
OSRD. 

Originally NDRC administered its work through five 
divisions, each headed by one of the NDRC members. 

These were: 

Division A — Armor and Ordnance 
Division B — Bombs, Fuels, Gases, & Chemical Prob- 
lems 

Division C — Communication and Transportation 
Division D — Detection, Controls, and Instruments 
Division E — Patents and Inventions 


In a reorganization in the fall of 1942, twenty-three 
administrative divisions, panels, or committees were cre- 
ated, each with a chief selected on the basis of his out- 
standing work in the particular field. The NDRC mem- 
bers then became a reviewing and advisory group to the 
Director of OSRD. The final organization was as follows : 

Division 1 — Ballistic Research 

Division 2 — Effects of Impact and Explosion 

Division 3 — Rocket Ordnance 

Division 4 — Ordnance Accessories 

Division 5 — New Missiles 

Division 6 — Sub-Surface Warfare 

Division 7 — Fire Control 

Division 8 — Explosives 

Division 9 — Chemistry 

Division 10 — Absorbents and Aerosols 

Division 11 — Chemical Engineering 

Division 12 — Transportation 

Division 13 — Electrical Communication 

Division 14 — Radar 

Division 15 — Radio Coordination 

Division 16 — Optics and Camouflage 

Division 17 — Physics 

Division 18 — War Metallurgy 

Division 19 — Miscellaneous 

Applied Mathematics Panel 

Applied Psychology Panel 

Committee on Propagation 

Tropical Deterioration Administrative Committee 


IV 


DECLASSIFIED 
By authority Secretary of 


NDRC FOREWORD 


MARKING S Must BE~CANCgTXynT 


AUG 2 6 1960 

As events of the years preceding 1940 revealed 
A ^fiftusness of the 

world situation, manjr -s cientists in this country 
came t(j ^ i$fo&R¥e O^dC^^^gBtBStfg scientific re- 
search for service in a national emergency. Recom- 
mendations which they made to the White House 
were given careful and sympathetic attention, and 
as a result the National Defense Research Commit- 
tee [NDRC] was formed by Executive Order of the 
President in the summer of 1940. The members of 
NDRC, appointed by the President, were instructed 
to supplement the work of the Army and the Navy 
in the development of the instrumentalities of war. 
A year later, upon the establishment of the Office 
of Scientific Research and Development [OSRD], 
NDRC became one of its units. 

The Summary Technical Report of NDRC is a 
conscientious effort on the part of NDRC to sum- 
marize and evaluate its work and to present it in a 
useful and permanent form. It comprises some 
seventy volumes broken into groups corresponding 
to the NDRC Divisions, Panels, and Committees. 

The Summary Technical Report of each Division, 
Panel, or Committee is an integral survey of the 
work of that group. The first volume of each group’s 
report contains a summary of the report, stating the 
problems presented and the philosophy of attacking 
them, and summarizing the results of the research, 
development, and training activities undertaken. 
Some volumes may be “state of the art” treatises 
covering subjects to which various research groups 
have contributed information . Others may contain 
descriptions of devices developed in the laboratories. 
A master index of all these divisional, panel, and 
committee reports which together constitute the 
Summary Technical Report of NDRC is contained 
in a separate volume, which also includes the index 
of a microfilm record of pertinent technical labora- 
tory reports and reference material. 

Some of the NDRC-sponsored researches which 
had been declassified by the end of 1945 were of 
sufficient popular interest that it was found desir- 
able to report them in the form of monographs, such 
as the series on radar by Division 14 and the mono- 
graph on sampling inspection by the Applied Mathe- 
matics Panel. Since the material treated in them 
is not duplicated in the Summary Technical Report 
of NDRC, the monographs are an important part 
of the story of these aspects of NDRC research. 

In contrast to the information on radar, which is 


of widespread interest and much of which is released 
to the public, the research on subsurface warfare is 
largely classified and is of general interest to a 
more restricted group. As a consequence, the report 
of Division 6 is found almost entirely in its Sum- 
mary Technical Report which runs to over twenty 
volumes. The extent of the work of a division can- 
not therefore be judged solely by the number of 
volumes devoted to it in the Summary Technical 
Report of NDRC: account must be taken of the 
monographs and available reports published else- 
where . 

The beginning of World War II found the United 
States with no program for the development of 
rocket weapons . By the end of the war this country 
was well in the lead, thanks largely to the efforts of 
Division 3. As a result of proposals by Dr. C. N. 
Hickman, NDRC rocket work was initiated in 1940 
under Division A, with Richard C. Tolman as chair- 
man. The work was carried forward by Division 3 
under two chiefs, John T. Tate in 1943 and Fred- 
erick L. Hovde through 1945. 

The program, carried out by several contractors 
with Army and Navy cooperation, produced rockets 
used effectively by our Infantry, Artillery, Navy, 
and Air Forces against submarines, ships, tanks, 
beach defenses, and inland positions. By virtue of 
their lack of recoil, rockets could be launched from 
men’s shoulders, automotive vehicles, small and 
large ships, and aircraft. One of the first to go into 
combat was the bazooka, the Infantry’s famed 
Panzer destroyer. In landing operations the Navy 
used barrage rockets effectively to smother Japa- 
nese shore defenses. From one Division 3 contract 
came also important contributions to the develop- 
ment of torpedoes and depth bombs. 

The Division 3 Summary Technical Report, pre- 
pared under the direction of the Division Chief and 
authorized by him for publication, outlines the 
technical and military knowledge resulting from 
this program. The performance of Division 3 in 
discovering and summarizing this information, and, 
even more, in applying it in timely development of 
new rocket weapons, deserves our admiration and 
gratitude. 

Vannevar Bush, Director 
Office of Scientific Research and Development 

J. B. Conant, Chairman 
National Defense Research Committee 


v 





























































































































FOREWORD 


D ivision 3 directed its operations toward two 
principal, and conflicting, objectives. The 
first was to develop rocket ordnance which the 
Army and Navy could and would use as early as 
possible in World War II. The second was to 
provide the new knowledge necessary as a basis 
for development of improved designs and addi- 
tional types of rocket weapons during a war of 
uncertain length. Maintaining the proper bal- 
ance between these aims as the war progressed 
was a matter of some difficulty, and was achieved 
only imperfectly. 

Most of the Division 3 rockets were developed 
to provide our military and naval forces with 
added fire and bombing power to meet tactical 
situations for which conventional artillery and 
bombs were unsuited or not effective. Except 
for the 1200-pound “Tiny Tim” aircraft rocket, 
all were under 200 pounds in weight. And none 
of the artillery type service rockets exceeded 
1600 feet per second in velocity. All of them 
employed grains of solid double-base propel- 
lants. None had wings or controls. 

Within these general limits, the work of Divi- 
sion 3 embraced research, development, design, 
experimental and pilot production, and many 
kinds of testing. Certain studies and develop- 
ments in underwater ordnance were carried on 
in close association with the broader activities of 
Division 6 in this field. 

In 1940 neither the Army nor the Navy had 
any rocket projectiles in service or under devel- 
opment. In the period 1942-1945 many types 
and sizes of rockets, components, launchers, and 
related ordnance items developed entirely or in 
part by Division 3 were used in combat in signifi- 
cant quantities and with substantial effects. 
Among these were the “bazooka” rocket, the 
“mousetrap” antisubmarine rocket, several 
types of rockets used primarily for barrages in 
landing and field artillery operations, and a vari- 
ety of rockets for aircraft armament. In addi- 
tion, the division’s laboratories doubled the 
range of the conventional 4.2-inch mortar 
through the development of new powder 
charges; another project involved structural 
modifications of the Mark 13 aircraft torpedo 


which increased the overall effectiveness of this 
important weapon several-fold. 

In or on the verge of production when the 
Japanese surrendered were “superbazooka” 
rockets, a recoilless 4.2-inch rifle, smokeless 
rockets for assisting the take-off of airplanes 
and flying boats, and numerous improved types 
of rocket ordnance already in service. Among 
the items in advanced development were water- 
discriminating fuzes for rockets fired from air- 
craft against ships, powder-powered launchers 
for V-l type flying bombs, powder-pressurized 
flame throwers, rocket propulsion units for mine 
field clearance devices, and proximity-fuzed 
rockets for defense against suicide aircraft at- 
tacks. 

In connection with these developments the 
division workers mastered many techniques and 
amassed much knowledge of rockets and other 
ordnance. This book provides a partial sum- 
mary of that knowledge, and a guide to much 
more of it. Not the least of the division’s accom- 
plishments has been the production and wide 
distribution of a large volume of reports on its 
work. 

Throughout its life the division provided con- 
sulting and other technical services to both Army 
and Navy, not merely on their own developments 
and those of the division, but also in connection 
with intelligence covering energy develop- 
ments. Field technical assistance was provided 
in the Pacific, in Great Britain, and in France. 
Of continuing value to the Army and Navy are 
the personnel and facilities acquired by transfer 
in the process of demobilizing Division 3. Many 
of the division’s principal operations are con- 
tinuing under the Navy Bureau of Ordnance. 

This book was prepared primarily for the use 
of military personnel entering on duties involv- 
ing research and development of rockets and un- 
derwater ordnance, technically competent in 
ordnance engineering, but with limited knowl- 
edge of these particular fields. The aims have 
been to summarize the “state of the art” as it 
developed during the war, and to indicate some 
of the directions of future research and develop- 
ment which appeared to be most promising or 


vii 


viii 


FOREWORD 


most necessary. The book serves as an introduc- 
tion to the numerous final and other technical 
reports submitted by the several division con- 
tractors. 

In scope this Summary Technical Report does 
not cover completely the activities of the divi- 
sion. The book is devoted primarily to basic 
phenomena, analysis, and methods ; the develop- 
ment and design of weapons and other equip- 
ment is covered only generally. In Chapters 18, 
19, and 20, C. W. Snyder sketches the evolution 
of most of the rocket designs developed under 
Section L. It is regrettable that there is no com- 
parable survey of the numerous items developed 
under Section H ; however, complete reports on 
these have been distributed. Fuzes, launchers, 
and rocket heads are treated only briefly. Among 
the subjects not covered at all are production, 
fire control, terminal ballistics, and tactical 
employment. Army and Navy experience in 
rocket development, production, testing, train- 
ing, and combat employment is not included, ex- 
cept indirectly as it affected the work of the 
division. The book is historical only where such 
treatment seemed to its authors to give the most 
effective exposition. 

In Part I, Dr. Max Mason and Dr. F. C. Lind- 
vall summarize the underwater ordnance activi- 
ties carried out in Division 3 to supplement the 
broader program of Division 6. Dr. B. H. Sage, 
in Part II, and Dr. R. E. Gibson, in Part III, treat 
the problems which lie at the core of rocket 
development, namely, those of propellants and 
interior ballistics. C. W. Snyder covers complete 
rockets, their launchers and their uses in Part 
IV, and the theories underlying their design and 
performance in Part V. 

Other volumes of the NDRC Summary Tech- 
nical Report Series include subjects related to 
the work of Division 3, as follows : 

Division 1 Propellants, interior ballistics, 

gun erosion 

Division 2 Terminal ballistics, choice of 

weapons (including rockets) 
for specified targets 

Division 4 Proximity and other fuzes for 

rockets, “tossing” of rockets 
from airplanes 

Division 6 Antisubmarine weapons, air- 

craft torpedoes, hydrodynamics 


Division 

7 

Fire control for rockets 

Division 

8 

Propellants, long-burning rock- 
ets, and high explosives 

Division 

11 

Flame throwers and incendiary 
rockets 

Division 

12 

Use of barrage rockets from 
DUKW’s 

Division 

14 

Radar ranging for aircraft 
rocket fire control 

Division 

18 

Metallurgy applicable to rock- 
ets 

Division 

19 

Rocket armament for guerilla 
warfare 

Applied 


Theory of heat transfer and of 

Mathematics 

nozzles, analysis of propellant 

Panel 


specifications 


The NDRC rocket development program was 
initiated in 1940. Its foundations were laid in 
Division A under the wise and far-sighted guid- 
ance of its Chairman, Dr. Richard C. Tolman, its 
Vice-Chairman, Dr. Charles C. Lauritsen, the 
Chairman of its Section H, Dr. Clarence N. 
Hickman, and, in 1942, the Chairman of its Sec- 
tion C, Dr. John T. Tate. In the NDRC reor- 
ganization at the end of 1943 these two sections 
were merged to form Division 3, with Doctor 
Tate as Chief. The program continued to grow 
rapidly. In the summer of 1943 Doctor Tate re- 
signed to devote full time to his responsibilities 
as Chief of Division 6. 

In September 1943, I became Chief of the 
Division and Acting Chief of its Section L, which 
was, in effect, a re-established Section C. Sec- 
tion H was reconstituted with Doctor Hickman 
as Chief. This organization continued through 
1945. Principal personnel of these several or- 
ganizations is shown in an appendix. 

The experience of Division 3 demonstrates 
conclusively that nonmilitary scientists can 
grasp quickly the needs of the fighting arms and 
the problems of the supply services, develop 
new and improved weapons and equipment rap- 
idly, within the limitations of available knowl- 
edge, expand that knowledge as required, and on 
this basis develop still newer and better items. 
In initiating such a program on the eve of war, 
the principle of exploring thoroughly, yet 
quickly, and correlating the technical knowl- 
edge available with the apparent operational 
needs of the war requires no defense. The impor- 


FOREWORD 


ix 


tance of bringing the best scientists into the pro- 
gram as early and in as large numbers as pos- 
sible has been proved; only thus can effective 
leadership be provided. Facilities must be pro- 
vided rapidly, but with a view toward expansion 
by severalfold. Constant evaluation of promise, 
progress and results is called for, as a basis for 
any needed redirection. 

It became apparent that the military principle 
of economy of force applies perhaps more 
strongly to wartime research and development. 
This is to say, more valuable results can be 
achieved sooner by early concentration on those 
few objectives of greatest value or promising of 
earliest attainment, to the exclusion, at least 
temporarily, of perhaps more attractive but less 
significant objectives. However, small holding 
and scouting forces are always needed, to con- 
solidate developments and to discover other 
promising lines of attack. The experience of 
the division showed the values of follow-through 
by the applied science forces into the fields of 
production, testing, training, and analysis of 
performance under conditions of ultimate serv- 
ice. Another analogy with military operations 
became apparent, namely, the necessity for 
prompt and complete abandonment of certain 
projects as soon as there is a conclusive deter- 
mination that, in comparison with other proj- 
ects, the probabilities of early enough success 
are not in proportion to the effort required. Fi- 
nally, the experiences of this division and others 
established new highs in teamwork between mil- 
itary personnel and scientists outside of the mili- 
tary organizations. 

Under the present conditions of peace, with 
time scale and other factors radically changed, 
research and development operations by or for 
the services must be governed by principles dif- 
fering somewhat from those above. I am con- 
vinced that the services must continue to have 
principal responsibility for the development of 
new weapons and other instrumentalities of 
warfare. Further, the services must provide for 
and supervise much more applied research, es- 
pecially in the fields of their specialized require- 
ments, than heretofore. For many reasons it 
seems both wise and necessary that they con- 
tinue strong fundamental research activities in 
their own military laboratories, yet at the same 


time promote an extensive and thorough extra- 
mural research program in order that the civil- 
ian scientists of the nation may continue to serve 
the needs of national defense in peace as well as 
in war. 

Whatever success the division attained is due 
in large measure to Dr. Vannevar Bush, Direc- 
tor of the Office of Scientific Research and Devel- 
opment, and to Dr. Irvin Stewart, Executive 
Secretary and Contracting Officer, and their 
staffs. Under their wise policies, flexible organ- 
izations and effective operating procedures, a 
majority of the nation’s scientists and scientific 
organizations performed an unprecedented job 
with a degree of efficiency and coordination un- 
usual in government operations in war or in 
peace. A basic element was the freedom allowed 
the divisions and contractors in choosing and 
using various means for achieving approved ob- 
jectives. Dr. James B. Conant, Chairman, and 
the members of the National Defense Research 
Committee, with their staffs, were responsible 
for approving the proposals of Division 3, and 
for reviewing and coordinating its work with 
that of other divisions. 

To the British government and to British sci- 
entists we owe a tremendous debt for making 
freely available their knowledge and experience 
gained in several years of defense research prior 
to the advent of NDRC and in active warfare 
preceding that of the United States. On the 
OSRD Liaison Office fell the burden of arrang- 
ing for and handling this international ex- 
change of information and of scientific per- 
sonnel. This exchange, especially in the early 
years, made possible a manifold increase in the 
division’s rate of progress. 

Liaison organizations and offices of the War 
and Navy Departments, and their cooperating 
field units, provided guidance as to specific serv- 
ice needs, participated in some phases of Divi- 
sion 3 developments, and expedited their transi- 
tions to combat employment. 

Many other NDRC divisions made available 
knowledge and services to hasten Division 3 
work, and included in their programs comple- 
mentary projects which increased the utility of 
Division 3 developments to the Armed Forces. 

The functions of initiating, establishing, 
guiding, supervising, and administering the op- 


X 


FOREWORD 


erations of Division 3 were well performed by 
its highly competent members, consultants and 
staff, and by the able staffs of the two sections. 
I am deeply grateful to all of them for faithful 
and talented services and for the privilege of 
working with them. 

The principal credit, of course, must go to the 
several contracting organizations (listed in an 
appendix) under which all of the Division 3 re- 
search and development was carried out. To 
them, and even more to their personnel, who fur- 
nished the ideas, knowledge, skills, and plain 
hard work which constituted the program, is due 
whatever praise the division may have earned. 

In conclusion, I express my appreciation to 


the six authors who contributed to this Sum- 
mary Technical Report. For it they gave of 
their time, talents, and efforts in the face of the 
pressing demands of their postwar activities, 
with little indication that the results would be 
worth the effort. As for myself, I am confident 
that they have produced a volume which will 
provide proper perspective for the numerous re- 
ports of the division, and which will, in conjunc- 
tion with those reports, preserve most of the 
benefits of the division’s five years of wartime 
ordnance development. 

Frederick L. Hovde 
Chief, Division 3 


O 


PREFACE 


T he general scope and results of the Division 3 
program are indicated in the Foreword by Fred- 
erick L. Hovde. The activities of the Division in- 
volved the services of approximately 800 scientists 
and engineers working under eleven prime con- 
tracts during the period 1940-1945. Total costs 
were of the order of $25,000,000 for research and 
development and $50,000,000 for experimental and 
pilot production. 

As a part of the effort to preserve the values of 
the Division’s work, this summary technical report 
was prepared, primarily for the orientation of tech- 
nical officers, engineers, and scientists who seek to 
acquire familiarity with the basic phenomena of 
solid fuel rockets or of the entrance of underwater 
ordnance into water. The volume may be useful 
also to more experienced workers in these fields, 
for review or reference purposes. The principles 
and important results of the Division program are 
summarized as of the end of 1945, as a foundation 
for the study of the substantial advances made 
thereafter by others. 

In this summary, the treatment of the subjects 
listed in the Contents, though it is technical, does 
not require previous knowledge of the subjects. 
Throughout the book, the emphasis is on technical 
considerations pertinent to military applications. 
Chapter 14 includes analyses of the military utility 
of solid fuel rockets. 

The information in this report is arranged in 
five parts by authors and subjects, rather than by 
projects. Each chapter was written by a single 
author who led Division 3 developments in the 
fields which he treats. The four authors of Parts I, 
II, IV, and Y were associated with the single Sec- 
tion L contract, number OEMsr-418 with the Cali- 
fornia Institute of Technology. The two authors 
of Part III were concerned with the activities under 
all ten Section H contracts; they were affiliated 
with the Allegany Ballistics Laboratory, which 
was operated by the George Washington Univer- 
sity. The fact that each of the six authors has 
written mainly on the experience in his organiza- 
tion, and in a manner of his own choosing, has 
resulted in a division of the text of this report on 
the basis of the sections and contracts indicated. 

As a result of this situation, the very important 
subjects of propellants and interior ballistics are 


presented from three points of view. In Part II 
Dr. Sage analyzes the problems of developing, de- 
signing, and producing rocket propellant charges 
of compositions of the sort employed in all United 
States rockets which saw combat in World War II. 
These compositions are generally similar to that of 
trench mortar sheet powder. In Part V, C. W. 
Snyder reviews these problems from the viewpoint 
of the projectile designer. Dr. Gibson and Dr. 
McClure describe in Part III the behavior of solid 
propellants of a much broader range of chemical 
composition. 

The functions of the volume technical editor 
have varied for different parts of the report, but 
in general they have been limited to minor revi- 
sions and rearrangements of the authors’ material, 
and the addition of somewhat inadequate footnotes, 
most of them referring to related subject coverage 
by the other authors. 

Mathematical treatments have been limited to 
relationships of fundamental importance, with de- 
tails of their derivation and application covered 
only by references to other reports. The mathe- 
matical symbols are consistent for each author but 
not entirely uniform among them. Most of the sym- 
bols are the same as those used in reports previously 
issued by the authors’ organizations. 

Because of the pressure of more urgent work, it 
was not possible to start the writing of this sum- 
mary technical report before the surrender of 
Japan. After that, progress on it was delayed by 
the discharge of the authors’ responsibilities in 
connection with final reports, contract termina- 
tions, transfer of many activities and facilities to 
the Services, and postwar engagements. Under 
these and other difficulties the six authors labored 
manfully to produce the following report. It is the 
editor’s opinion that thfe advantages derived from 
their superior qualifications in the subjects covered 
have amply justified the acceptance of the delays. 
The authors and the editor have reviewed the 
galley proofs, but the tight publication schedule 
has precluded this process on the page proofs. 

This volume is a somewhat incomplete summary 
of the scientific and technological results of Di- 
vision 3 work. It was not possible, unfortunately, 
to include much information on the rocket pro- 
jectiles developed under Section H, or on the nu- 


xi 


xii 


PREFACE 


merous applications of rocket technology by that 
section to the development of rocket thrust units 
for airplanes and anti-mine devices, of recoilless 
guns, and of devices which utilized the burning of 
rocket propellants as sources of high pressure gases 
for several purposes. This report outlines the basic 
principles. For complete information on these and 
other Division 3 developments, the reader is re- 
ferred to the General Bibliography appended, in 
which are listed several hundred of the more im- 
portant reports of the Division. 

In keeping with its character as a technical sum- 
mary, this report includes information on Division 
3 personnel, organization, contracts, and projects 
only as listings in appendices. No attempt has been 
made to present the history of rockets or of the 
Division’s work on them, or to describe the combat 
or other Service experience with Division 3 devel- 
opments. 

A popular account along these lines is available 
from the Superintendent of Documents under the 
title “Rocket Ordnance — Development and Use in 
World War II.” Little, Brown and Company have 
published a series of volumes on OSRD and its 
contributions to World War II. Of these, the one 
by Dr. James P. Baxter 3rd is the short history 
of OSRD. Of the other long history volumes, about 
half of the one edited by Professor John E. Bur- 
chard is a history of Division 3 work, another by 
Burchard and Thiesmeyer describes the work of 
OSRD scientists, including several from Division 
3, in combat areas, and another, by Dr. Irvin 


Stewart, outlines the organization and adminis- 
tration of OSRD. 

For many reasons, this report has excluded 
acknowledgments of credit for technical or other 
contributions to the advancement of the Division 
program. The titles of reports listed in the appended 
General Bibliography provide some indications as 
to the types of contributions made by their authors. 
The work of the Division was aided greatly by 
lessons learned from the experience of United States 
and British Service and civilian agencies in the de- 
velopment, production, testing, and training and 
combat use of rockets and other ordnance. 

The editor acknowledges his gratitude to all the 
authors for the cooperation they provided under 
difficult conditions in the preparation of this re- 
port. It is hoped that the readers will find enough 
value in their chapters to justify a generous toler- 
ance of editorial defects. Dr. Gibson, Dr. McClure, 
and the editor join in acknowledging the helpful 
review and comment on Part III provided by Dr. 
Alexander Kossiakoff, former Deputy Director of 
the Allegany Ballistics Laboratory. Taking ad- 
vantage of this opportunity, the editor records here 
the great satisfaction he has derived from several 
years of pleasant associations with the personnel 
of OSRD, NDRC, and many of their contracting 
organizations, and in particular with Dr. Richard 
C. Tolman, Dr. John T. Tate, and Frederick L. 
Hovde. 

Eliot B . Bradford 
Editor 


CONTENTS 


PART 1 

UNDERWATER ORDNANCE 
by E. B. Bradford 

CHAPTER PAGE 

1 Antisubmarine Weapons and Underwater Ballistics 

by Max Mason 3 

2 Aircraft Torpedo Development and Testing by F. C. 

Lindvall 13 

3 Basic Research on Torpedo Entrance Phenomena by 

F. C. Lindvall 16 

4 Facilities and Instrumentation for Study of Torpedo 

Entry by F. C. Lindvall 21 

PART II 

ROCKET PROPELLANTS AND INTERIOR BALLISTICS 
by B. H. Sage 

5 Interior Ballistics by B. H. Sage 39 

6 Ignition by B. H. Sage 52 

7 Dry-Processed Double-Base Propellants by B. H. Sage . 56 

PART 111 

ROCKET ORDNANCE: THERMODYNAMICS AND 
RELATED PROBLEMS 
by R. E. Gibson 

8 Types of Rocket Propellants by R. E. Gibson ... 67 

9 Thermodynamic Problems by F. T. McClure ... 71 

10 Kinetic Problems by R. E. Gibson 78 

11 Structural Problems by R. E. Gibson 89 

12 Interior Ballistics Problems by F. T. McClure ... 96 

13 Properties of Rocket Propellants Available or Devel- 
oped during World War II by R. E. Gibson .... 99 


xiv 


CONTENTS 


PART IV 

ROCKET WEAPONS AS DEVELOPED AND USED IN 


WORLD WAR II 
by C. W. Snyder 

CHAPTER PAGE 

14 Military Needs Which Rockets Can Meet by C. W. 

Snyder 117 

15 Rocket Heads by C . W. Snyder 126 

16 Rocket Fuzes by C . W. Snyder 129 

17 Rocket Launchers by C. W. Snyder 138 

18 Service Designs of Fin-Stabilized Rockets for Surface 

Warfare by C . W. Snyder 148 

19 Service Designs of Fin-Stabilized Rockets for Aircraft 

Armament by C . W. Snyder 165 

20 Service Designs of Spin-Stabilized Rockets by C. W. 

Snyder 196 


PART V 

ROCKET ORDNANCE: THEORY , PRINCIPLES , AND 


DESIGN 
by E. B. Bradford 

21 General Theory of Rocket Performance by C. W. Snyder 211 

22 Design of Rocket Propellant Charges by C. W . Snyder . 223 

23 Motor Design by C. W. Snyder 244 

24 Exterior Ballistics of Fin-Stabilized Rockets by C . W. 

Snyder 267 

25 Exterior Ballistics of Spin-Stabilized Rockets by C. W. 

Snyder 288 

Bibliography 307 

OSRD Appointees 366 

Contract Numbers 368 

Service Project Numbers 370 

Index 373 


Jj 


SUMMARY 

by E. B. Bradford 


Underwater Ordnance 

Part I of this report describes briefly the un- 
precedented facilities developed at Morris Dam 
(near Pasadena) for full-scale studies of the be- 
havior of aircraft torpedoes and other under- 
water ordnance items on entry into water at 
extreme speeds and angles. With these and other 
facilities, important contributions were made to 
several of the weapons of World War II, and to 
better understanding of the phenomena of 
water entry and underwater travel. The Navy 
continued these operations after the war. High- 
lights of the wartime work are summarized in 
the Introduction to Part I. 


Solid Fuel Rockets 

Parts II-V summarize most of the principles 
and practices employed by Division 3 in the 
development of nearly all the rockets used by 
United States forces in World War II combat, 
and of several others not so used. In all these 
rockets smokeless powders were used. By the 
end of the war, several types of rockets had 
demonstrated their utility in many tactical situ- 
ations, and Navy procurement of them was on a 
financial scale comparable with conventional 
ammunition. 

Rocket Characteristics and Uses 

In nearly all their uses, rockets performed 
the function of artillery. Lethal or other pay- 
loads up to 500 lb were delivered to ranges up to 
10,000 yd, with detonation or other effects. By 
virtue of their self-contained recoil-less propul- 
sion, and the light, simple launchers thus made 
possible, rockets achieved big-gun effects from 
such relatively frail mounts as airplanes, small 
boats, light land vehicles, and men’s shoulders. 
Fired forward from airplanes, fin-stabilized 
rockets in calibers up to 12 in. were especially 
useful against small hard targets. For faster 
airplanes, spin-stabilized rockets offer certain 


advantages. Rockets used from surface ships 
included the “mousetrap” antisubmarine depth 
bomb, several types (finners and spinners) for 
offshore barrages, and fast spinners (1,540 
ft/sec) as main batteries for PT boats. In 
ground warfare, rocket launchers mounted on 
trucks and tanks drenched area targets at cri- 
tical periods. 

The launcher plays no part in propulsion and 
is subjected to little or no recoil force. Its func- 
tion is simply to guide the initial motion of the 
rocket along the line of proper train and eleva- 
tion. This is accomplished by light rails, tubes 
or slots, or, on airplanes, by the airstream. 

On the other side of the picture it must be 
noted that rockets have disadvantages which 
may include rearward blast, smoke, flash, lack 
of accuracy, limited velocity and range, low per- 
centage of weight effective at the target, and 
variation of performance and safety with tem- 
perature. 

Rocket Heads, Fuzes, and Effects 

The effects achieved at the target by most 
rockets are those of artillery and aerial bombs. 
In elementary rocket theory the head is the 
first item selected or designed, on the basis of 
target effects desired. Since the accelerations 
and stresses of projection are low, the problems 
of head design are generally similar to those of 
bomb design. 

An advantageous property of long-finned 
rockets is their long straight underwater travel. 
This characteristic was improved, by blunting 
the nose curvature, so that 3.5-in. aircraft 
rockets with solid heads were enabled to per- 
forate submarines after 130 ft of underwater 
travel, thus making range estimation less criti- 
cal. 

The requirements of function and safety for 
rocket fuzes are the same as those for shell and 
bombs. Shell fuzes were adapted for spin- 
stabilized rockets. For fin-stabilized rounds, 
with no spin and low setback, the fuzes involved 
various combinations of mortar fuze adapta- 


SUMMARY 


tions, setback devices, arming wires, air-driven 
propellers, and time delay. Impact was usually 
used to trigger detonation, in some cases with 
time delay. 

An extensive series of fuzes was developed, 
of which many were standardized. One of the 
last fuze developments provides radically new 
performance, especially for underwater hits on 
floating targets. This deceleration discriminat- 
ing fuze arms partially on first impact with 
water or target but fires only after it has pene- 
trated the hull (high deceleration) and emerged 
inside (low deceleration) or after its velocity 
has dropped to a low value. 

Exterior Ballistics 

The behavior of rockets in flight and the 
methods used for its analysis have many simi- 
larities to those of shell and bombs. The out- 
standing differences are due to the continuation 
of propulsion and acceleration over distances as 
much as 1,000 ft beyond the launcher. With 
spin-stabilized rockets the rate of spin con- 
tinues to increase throughout the period of pro- 
pulsion. Most of the dispersion of rockets has 
its origin in this period. 

Accuracy has been improved, and the factors 
affecting it have become better understood, as a 
result of thorough analyses of the oscillations, 
precessions, and nutations of rockets in flight. 
The flight behavior and especially the accuracy 
of World War II rockets were undesirably sen- 
sitive to changes in temperature. As indicated 
below, propellant developments late in the war 
improved this situation. Wind is a factor with 
several effects on rocket flight, some of them re- 
lated to temperature and all of them tending to 
reduce accuracy. 

Fin stabilization provides simplicity, econ- 
omy, flexibility in design, and possibilities for 
various combinations of a few heads and motors 
to serve many purposes. Spin stabilization has 
advantages in better accuracy, shorter launch- 
ers, easier handling and better adaptability to 
automatic launchers, but it introduces severe 
design restrictions. The requirements for flight 
stability involve relationships among velocity, 
rate of spin, propellant strength, ratio of length 
to caliber (commonly 6 to 7) and weight distri- 


bution. One result is that different types of use 
usually require different designs. 

Rocket Motors 

The function of a rocket motor is to provide 
an impulse for the acceleration of a projectile 
or other load. This total impulse is the product 
of the thrust and its duration, usually expressed 
in pounds-seconds. The rocket motor produces 
the thrust as a reaction to its rapid rearward 
discharge of a stream of gases. In the case of 
free flight, the impulse given to the whole rocket 
is equal to the momentum (mass X velocity) 
imparted to it, which is equal and opposite to 
the momentum given the gases. 

For each size and type of rocket there is an 
upper limit to the velocity obtainable, even with 
the payload reduced to zero. This limit can be 
raised by increasing the impulse-to-weight ratio 
of the motor, the motor specific impulse, com- 
monly expressed in pounds-seconds thrust per 
pound of initial weight of the loaded motor. 
This ratio is increased by designing for com- 
bustion at constant, low pressure in a chamber 
of high strength-to-weight ratio. A basic re- 
quirement is a propellant composition which, 
burned in a suitably designed motor, gives a 
high specific impulse. A value typical of World 
War II rocket propellants is 200 lb-sec thrust 
per pound of propellant burned. Multiplication 
of specific impulse by the acceleration of gravity 
gives the effective gas velocity, frequently used 
to indicate the performance of a propellant in 
a rocket. The velocity acquired by the rocket is 
roughly this effective gas velocity multiplied by 
the ratio of propellant weight to total weight. 

The typical solid fuel rocket motor is a steel 
tube, closed at the front end, with one or several 
venturi nozzles at the rear. The nozzles serve 
to maintain the desired combustion pressure, to 
smooth and direct the discharge of propellant 
gases, and, by expanding them, to add about 
30 per cent of the total thrust. Motors for 
finners are usually long and slim, for reasons of 
aerodynamics, accuracy and economy; spinner 
motors are rather short, as required for flight 
stability. Spin is produced by multiple nozzles 
mounted on a circle at angles resulting in a 
peripheral component of thrust. 


SUMMARY 


Charge Design 

Within the motor is the propellant charge, of 
weight given by dividing the specific impulse 
characteristic of its composition into the total 
impulse required. Constant pressure operation 
of the rocket motor requires a constant mass 
rate of discharge of propellant equalled by a 
constant mass rate of burning, the latter involv- 
ing parallel layer burning over a constant total 
surface which recedes at a constant linear rate 
of burning. Constant burning area may be se- 
cured simply by grain shape, or it may involve 
“inhibiting” certain surfaces to prevent their 
burning. High density of loading is sought; this 
leads frequently to a single grain charge. Other 
considerations may require a multi-grain 
charge. Low operating pressure is secured by 
a wide nozzle opening, a small burning area, 
and a propellant composition of slow linear 
burning rate. 

Characteristics of Solid Propellants 

Of fundamental importance in the interior 
ballistics of rockets are the linear burning rate 
of the propellant and the increase of this rate 
with pressure and with temperature. For the 
propellants used in the rockets which saw com- 
bat, the pressure sensitivity was such that the 
equilibrium motor pressure varied approxi- 
mately as the fourth power of several motor 


parameters ; newer propellants brought this 
power down to about 1.2. 

The temperature range within which the 
best World War II rockets gave safe and de- 
pendable performance was — 40 F to + 140 F. 
Pressure, thrust, acceleration, burning time, 
burning distance, and dispersion varied by fac- 
tors as high as three between the upper and 
lower limits, mainly because of the sensitivity 
of the burning rate to propellant temperature. 
Propellants developed during the war had tem- 
perature coefficients from 1.5 down to 0.1 per 
cent change in equilibrium motor pressure per 
degree centigrade. 

The improvements in propellant character- 
istics resulted from studied changes in chemical 
composition. The physical properties of pro- 
pellants, especially mechanical toughness, are 
important to proper performance under the 
stresses of rocket acceleration. The composi- 
tions and characteristics of solid rocket pro- 
pellants are surveyed in this report, as are proc- 
esses for propellant production. 

Conclusion 

Many possibilities for rockets substantially 
better than those of World War II have been 
demonstrated ; others are indicated. Several 
chapters of this report include recommenda- 
tions as to promising lines for future research 
and development. 





PART I 


UNDERWATER ORDNANCE 

By E. B. Bradford a 


I n its development of rocket ordnance Division 3 
and its predecessor units of NDRC led the way 
in virgin territory; in 1940 neither the Army nor 
Navy had any activities or much interest in this 
field. In underwater ordnance, on the other hand, 
the Services, especially the Navy, had extensive 
experience and activity. Nevertheless, the civilian 
and largely academic scientists of NDRC were able 
to grasp the outstanding problems and contribute 
effectively to many of them, in the improvement of 
old weapons like torpedoes, in the development of 
new ones like ahead-thrown depth bombs, and in 
the general advance of underwater ordnance re- 
search, development, and testing. 

In NDRC, Division 6 (formerly Section C4) pur- 
sued rather broad programs on underwater ord- 
nance. 15 Certain specialized work in this field was, 
however, carried out in Division 3, in substantial 
part for Division 6. All this Division 3 work was 
done by two special sections of the rocket contract 
(OEMsr-418) with the California Institute of Tech- 
nology [CIT]. Section IV was concerned mainly 
with water entry and underwater performance char- 
acteristics of depth bombs, depth charges, and 
similar ordnance; its activities included full-scale 
testing of service and experimental ordnance items, 
model scale studies, and associated theoretical re- 
search. Section VII was concerned entirely with 
aircraft torpedoes, primarily with the fundamental 
study of the behavior of torpedoes and their com- 
ponents on high-speed entry into water in full-scale 
tests. 

Although both groups had as their prime function 
the providing of test data and other information for 
application elsewhere to problems of ordnance de- 
sign, both participated directly in certain weapon 
developments which found significant service uses. 
Among those involving Section IV were ahead- 
thrown depth bombs of both the spigot-projected 
(Hedgehog) and rocket-propelled (Mousetrap) 
types, retro rocket depth bombs for the attack of 
submarines by MAD-equipped airplanes, and the 

a Volume editor. 

b See Division 6 Summary Technical Report. 


forward-firing aircraft rockets which were so effective 
against underwater targets as well as others. Sec- 
tion VII, starting with model test indications from 
a Division 6 program, developed the shroud ring 
modification for the tail of the Mk 13 torpedo, and, 
in the summer of 1944, provided the first 1,000 of 
these to go to combat areas. This modification 
eliminated the serious restrictions imposed on pilots 
by the older torpedoes; with the new ones they were 
enabled to release their torpedoes at any speeds of 
which their airplanes were capable, from higher 
altitudes, and still secure more hot, straight runs 
than they had formerly from lower and slower 
approaches, with their greater exposure to A A fire. 

In Chapter 1 of this Division 3 Summary Tech- 
nical Report, Dr. Max Mason, who headed Section 
IV, outlines its principal activities and results. In 
Chapters 2, 3, and 4, Dr. F. C. Lind vail summarizes 
the Section VII work under his supervision. Both of 
these summaries indicate the scopes of the programs 
and of the special facilities and instrumentation de- 
veloped for them. Each serves as an introduction 
to an OEMsr-418 final report volume (cited) on the 
work. Several hundred copies of each of these 
volumes have been distributed through the War 
and Navy Departments. 

Sections IV and VII were both set up initially 
to provide and operate new and unprecedented 
facilities for the securing of full-scale test data not 
obtainable as accurately or as economically by exist- 
ing practices. The principal facilities of both 
groups are located at the Morris Dam Reservoir in 
Southern California. Together with records and 
experienced personnel, they were taken over by the 
Navy in late 1945. They are now being expanded 
and operated under the Naval Ordnance Test Sta- 
tion, Inyokern, California, as parts of the Navy’s 
peacetime underwater ordnance program. The 
Section IV facilities were designed and used to pro- 
duce data with laboratory precision from full-scale 
launchings duplicating pertinent conditions of oper- 
ational use of several types of underwater ordnance. 
The data obtained covered air- water trajectories, 
accuracy, sinking speed, and fuze functioning, as 


1 


2 


UNDERWATER ORDNANCE 


well as the effects of shape and weight distribution 
on these aspects of performance. The Section VII 
facilities provided for the launching of torpedoes 
into water at extreme speeds and angles; rather 
elaborate external and internal instrumentation was 
employed to provide detailed information on the 
behavior of torpedoes and their components under 
these conditions. Thus, in both cases, it was 
possible to get more, and more accurate, informa- 
tion than that obtainable from service-type tests, 
with their complications as to time, weather, man- 
power, availability and limitations of airplanes, 
ships, equipment, etc. The method previously used 
for securing comparable data on torpedoes, for ex- 
ample, had been to drop them from available air- 
planes (frequently not fast enough) and try to see 
what happened — the limitations are obvious. With 
the new facilities, many features of underwater 
ordnance designs could be established more defi- 
nitely at earlier stages of development, with service- 
type testing required for little more than final proof. 

In both sections programs of basic research were 
carried on in association with the testing activities, 
to provide foundations for further advances in 
underwater ordnance. These programs are out- 
lined by Mason and Lind vail, and presented in 


detail in the CIT final reports which they cite as 
bibliographic references. 

To complete the picture of Division 3 torpedo 
work, an early, stopgap development may be men- 
tioned briefly. In 1943, in an effort to provide a 
way around the limitations of the Mk 13 torpedo, 
CIT developed a device which decelerated it by 100 
knots between release and water entry. This was 
accomplished by an assemblage of rocket motors so 
mounted on the torpedo as to exert rearward thrust 
during the free fall, and to detach itself before entry. 
Such devices performed successfully in torpedo- 
dropping tests at the San Diego Naval Air Station 
and the Newport Naval Torpedo Station, but were 
not adopted for service. 

In considering the summaries by Mason and Lind- 
vall, it must be remembered that World War II 
ended with the various research programs in widely 
differing stages of completion. Hence, although 
much has been learned about some items, there are 
many others in which the surface had barely been 
scratched by the time the activities under the 
OSRD contract were taken over by the Navy. In 
these cases the results should be considered as pre- 
liminary surveys indicative of the direction in which 
further work might fruitfully be pursued. 



Chapter 1 

ANTISUBMARINE WEAPONS AND UNDERWATER BALLISTICS 


By Max Mason a 


11 INTRODUCTION 

T he underwater ordnance studies of Section 
IV of the organization which grew up at the 
California Institute of Technology under Contract 
OEMsr-418 had two main aspects: (1) the building 
up of special facilities at Morris Dam, and their use 
in tests and development of antisubmarine ord- 
nance, and (2) mathematical and model scale 
studies of the fundamental ballistics of water entry 
and underwater travel. These are covered under 
separate headings in this chapter. 

12 FULL-SCALE WEAPON TESTING 
AND DEVELOPMENT 

Throughout World War II Morris Dam con- 
ducted full-scale and large-model tests for which no 
comparable facilities were available elsewhere in 
this country. As a part of the testing program 
about fifty different service devices of the United 
States and British Navies were studied, and meas- 
urements of their underwater performance reported b 
for evaluation, guidance of design changes, and 
other uses. 

Similar testing services were provided for Divi- 
sion 6 (formerly Section C4) . Among the ordnance 
items to which Morris Dam contributed in this way 
were the following: 

Depth charges, Mks VI, IX, XII, and XVII. 

U. S. versions of the British Hedgehog projectile. 
The Mk 24 mine. 

The antisubmarine scatter bomb of Divisions 3 
and 6. 

The British Projectile Type C (Squid). 

a Supervisor of Section IV (Underwater Properties of Pro- 
jectiles) of Contract OEMsr-418 at the California Institute of 
Technology. 

b All reports issued by Section IV are included in the 
general bibliography appended to this volume, under OEMsr- 
418 file series IBC, IEC, IHC, IIC, IOC, IPC, JHC, and JPC. 
The bibliography of Water Entry and Underwater Ballistics of 
Projectiles 1 lists these reports under several subject headings. 
They are listed also by a different subject classification in the 
NDRC Summary Technical Report Microfilm Index. 


Numerous pistols and fuzes for these and other 
weapons. 

In addition to providing these test services, the 
Morris Dam group participated directly in the de- 
velopment of several types of rocket ordnance for 
the attack of underwater targets, as indicated 
below. 


1,2,1 The Problem of Antisubmarine 
Ordnance 

In the period following the first World War the 
detection and location of submerged submarines by 
echo ranging (“sonar”) was highly developed. By 
this means both direction and range of a submarine 
could be determined from a single ship. The stand- 
ard depth charge remained, however, the only 
ordnance for attack. This was a very ineffective 
weapon. Among its shortcomings were slow sink- 
ing speed and rather erratic underwater trajectories. 
Although such depth charges could be thrown from 
large ships, they had to be dropped from small ones. 
In both cases the number which could be launched 
from one ship simultaneously or in quick succession 
was limited . Their fuzes functioned at preset depths , 
whether near the submarine or not. Sound contact 
with the submarine was frequently lost because of 
the maneuvering required for dropping the depth 
charges and the disturbances caused by their ex- 
plosions. Better antisubmarine ordnance, prefer- 
ably usable from small ships, was urgently needed. 
This view was emphasized by the results of British 
statistical studies of depth charge attacks. 

Attention was therefore directed to fast-sinking 
bombs fuzed to detonate only on contact with the 
submarine, and to the projection of a number of 
such bombs forward from the hunting ship during a 
sonar fix. In this way cat-and-mouse tactics could 
replace the blind-man’s-buff method of the depth 
charge. The effectiveness of this type of antisub- 
marine armament was indicated by British work on 
the development of the “Hedgehog,” first of the 
“ahead-thrown” weapons. This consisted of an 


3 


4 


ANTISUBMARINE WEAPONS AND UNDERWATER BALLISTICS 


array of spigot launchers, from which a substantial 
number of contact-fuzed depth bombs, each con- 
taining about 35 lb of high explosive, were pro- 
jected over the bow. In the absence of a hit, there 
was no explosion, and sound contact was retained. 

1,2,2 Establishment of Morris Dam 
Laboratory 

Development of “ahead-thrown” ordnance for 
U. S. production and use required facilities for 
studying behavior of the projectiles on entering into 
and proceeding under water, and for observing fuze 


Scientific Research and Development; technical 
supervision for the government was the responsibility 
of Section C4 (later Division 6) of NDRC. The 
engineering talent for design and operation of the 
new facilities came mainly from the CIT 200-in. 
telescope project, on which activity was suspended 
during World War II. A general view of the instal- 
lation is shown in Figure 1. The first work at the 
Morris Dam was in cooperation with other C4 
activities at New London, on the testing of depth 
charges and the design of fast-sinking bombs. 
Studies of rocket-propelled antisubmarine ordnance 
soon became an important activity, and from 
June of 1942 the activities were included in those 



Figure 1 . General view of Morris Dam and testing facilities. Splash near the center of the picture indicates 
projectile has just been launched down one of the ramps. Nets and targets used for determining trajectories 
and recovering the projectiles are shown at the right. 


action. To meet these and related needs, the under Contract OEMsr-418, which covered the CIT 
Morris Dam Laboratory was established in August rocket developments then under Section C of Divi- 
1941, by the California Institute of Technology sion A, NDRC, and, after December 1942, under 
under Contract OEMsr-329 with the Office of Division 3. 



FULL-SCALE WEAPON TESTING AND DEVELOPMENT 


5 


1 2,3 Mousetrap — an “Ahead-Tlirown” 
Depth Bomb without Recoil" 

Because of its recoil effects, the Hedgehog was 
usable only on fairly large ships, with well-braced 
foredecks. In 1942 there were not available enough 
such craft to meet the urgent submarine situation. 
To provide equivalent striking power for smaller 
craft, CIT developed a weapon similar in use and 
effectiveness to the Hedgehog, but with recoilless 
rocket projection instead of spigot gun projection. 

This armament, known as “Mousetrap,” resulted 
from collaboration of the Morris Dam group with 
the rocket group. Its development involved deter- 
mination of the best head shape, weight distribution, 
and fin configuration to provide maximum accuracy 
in launching, air flight, oblique water entry, and 
sinking. With this weapon many smaller ships were 
equipped with effectively the same attack power as 
destroyers, and antisubmarine patrols were sub- 
stantially strengthened . 

124 Retro Bombs for Antisubmarine 
Aircraft c 

The development of the magnetic airborne detector 
[MAD] presented an analogous problem. Until the 
advent of the sonobuoy, MAD was the only device 
by which an airplane could detect an invisible, sub- 
merged submarine. However, it indicated location 
only when directly over the submarine. Conven- 
tional aircraft armament was at a disadvantage in 
this situation. The rocket and underwater ordnance 
groups at CIT collaborated again, to conceive and 
develop a type of armament suitable for use with 
MAD. For the ammunition, heads adapted from 
the Mousetrap were used, mounted on rocket 
motors which propelled them at speeds to match 
aircraft cruising speeds. Usually mounted twelve 
under each wing, these were fired backward on 
MAD indications (after exploratory passes) to enter 
the water in a pattern across the area in which the 
submarine had been located. Here the problem 
was one of substantially vertical fall, water entry, 
and sinking, with the accuracy problem compli- 
cated by oscillation of the missiles at entry. 

c The Mousetrap rockets are described briefly in Chapter 18; 
retro-rockets and their components, launchers and employ- 
ment are covered at greater length in Bureau of Ordnance 
publications and other reports listed in the general bibliog- 
raphy appended to this volume. 


125 Aircraft Rockets for Underwater 

Targets" 

The third and most successful project on which 
the Morris Dam group collaborated with the Divi- 
sion 3 rocket workers at CIT was the development 
of rockets which, fired forward from diving aircraft 
to enter the water at high speed, and after some 
distance of underwater travel, would hit an under- 
water target with energy enough to penetrate the 
hulls of submarines and thin-skinned ships. Here 
again the ballistics of air flight, water entry, and 
underwater travel had to be combined to secure 
maximum range and accuracy and to determine the 
best dive angles (and hence water entry angles) for 
attacks. 

126 Facilities for Testing Underwater 

Performance 

A major part of the work of the Morris Dam 
group was the devising of instrumental means of 
study and measurement. The principal facilities, 
described in detail in the Section IV final report, 1 
are summarized in the following paragraphs. Ex- 
cept for item 6, all these are at Morris Dam. 

1 . A large sound range for observing time-position 
relations, with a horizontal recovery target 50 ft 
by 50 ft which can be lowered to 180-ft depth of 
water and a vertical target 62 ft by 70 ft for shallow 
entry. These can be seen in Figure 1. Continuous 
records are obtained from six hydrophones and a 
six-channel oscillograph. The coordinates of under- 
water trajectories are obtained without arithmetic 
reckoning by a special computing device. 

2. An electrical net, and other net equipment, for 
determining shallow trajectories which cannot be 
evaluated with sufficient precision by the sound 
range. 

3. Rocket and blowgun launching facilities, ad- 
justable for obtaining desired air trajectories or 
entry angles, with entry velocities as high as 1,000 
fps for 1-in. diameter specimens and about 900 fps 
with 70-lb projectiles. The high entry velocities are 
a special objective of this facility. 

4. Facilities for taking underwater motion pic- 
tures of bubble and cavitation phenomena down 
to the maximum depth of the lake. 

5. Facilities for underwater impact tests and fuze 
tests. 


6 


ANTISUBMARINE WEAPONS AND UNDERWATER BALLISTICS 



Figure 2. Behavior of unvented torpedo model. Side and bottom views of oblique entry show how the water 
clings to underside. 



FULL-SCALE WEAPON TESTING AND DEVELOPMENT 


7 



Figure 3. Behavior of model with vented nose. Side and bottom views of oblique entry show how venting 
has relieved under-pressure which produced turbulence shown in Figure 2. 



8 


ANTISUBMARINE WEAPONS AND UNDERWATER BALLISTICS 


6. In the laboratory at the Institute, a 24-ft by 
4-ft by 4-ft glass-walled model tank providing 
entry velocities up to 180 fps for 1-in. or 2-in. 
models. This is equipped with an entry whip re- 
corder, Edgerton-type stroboscopic lights, and a 
number of special types of high-speed cameras. 

13 MODEL SCALE AND THEORETICAL 
STUDIES OF WATER ENTRY BALLISTICS 

During the earlier part of World War II the 
demands on the Morris Dam facilities for study 
and test of service ordnance were so great that but 
slight attention could be paid to furthering the 


Later (in 1944) emphasis was placed on this type 
of work. The equipment mentioned in item 6 of 
the list of facilities was produced in the effort for 
quantitative results of precision on model behavior 
at water entry, including the effects of geometric 
scale, of entry velocity, and of other entry condi- 
tions upon the underwater behavior of projectiles. 
Some problems were attacked mathematically and 
checked by experiment. Extensive studies were 
made on models of the Mk 13 Mod 6 torpedo, to 
complement the full-scale tests carried out under 
Section YII of OEMsr-418. (See Chapter 2 of this 
volume.) 

The following paragraphs outline the scope of 
Water Entry and Underwater Ballistics of Projectiles / 



Figure 4. A shadowgraph, such as this one showing vertical entry of sphere, permits study of form of 
water surface during impact. 


understanding of the hydrodynamics of the water 
entry of projectiles. A small glass-walled tank with 
a launching catapult had been set up in the labora- 
tory and used for quantitative experiments with 
small models. Related studies of model behavior 
were conducted by the Alden Hydraulics Labora- 
tory at Worcester, Mass., under an OEMsr-418 
subcontract. 


the final report d under OEMsr-418 on the work of 
Section IV. 

An understanding of the details of dynamic be- 
havior of the entrance of projectiles into water must 
form the basis for the application of mathematical 
analysis to these complicated phenomena. These 

d Several hundred copies of this and other OEMsr-418 final 
reports were distributed to the Services. 


STUDIES OF WATER ENTRY BALLISTICS 


9 


details can be studied most accurately with models 
of greatly reduced scale by the aid of modern high- 
speed photography. A matter of primary im- 
portance for practical design is the ability to predict 


full-scale behavior from model behavior. The report 
devotes considerable length to this problem. An 
end result is that under-pressure in an air pocket on 
the lower surface of the nose of the projectile at 



Figure 5. Stroboscopic methods permit study of flow patterns of fluid in neighborhood of projectiles. Here, 
with 300 light flashes per second, displacements of illuminated bubbles show motion of water during vertical 
entry of steel sphere. 



10 


ANTISUBMARINE WEAPONS AND UNDERWATER BALLISTICS 


entry plays an important role in the behavior of 
small models; by relieving this underpressure by 
“venting/’ satisfactory modeling may result. The 
effect of such venting is illustrated in Figures 2 and 
3. The report also brings out the importance, in the 
case of high-velocity projectiles, of modeling on a 
velocity or “stress” basis, as contrasted with Froude 
scaling. 

Experimental results concerning the impact de- 
celeration of spheres and cones, the form of the 
water surface during impact, and the flow patterns 
of the fluid in the neighborhood of projectiles are 


Because phenomena associated with venting do 
not appear to have been discussed elsewhere, special 
attention is devoted to discussion of experimental 
and theoretical aspects of this important subject. 
Comparisons between model and prototype behavior 
have been presented in all cases for which adequate 
observational material was available. Although 
scale effects are apparent in the details of projectile 
behavior, high-velocity adequately vented models 
may, in general, be relied upon to reproduce the 
prototype trajectory within a few diameters over 
ten to twelve lengths of underwater travel. The 



Figure 6. Early development of separation film of air between solid and fluid. 


reported, together with tests establishing the ab- 
sence of appreciable tensional stress in the fluid 
during its separation from the projectile or of sig- 
nificant shear stress during impact. Figures 4 and 5 
illustrate photographic methods that were used in 
making these studies. 

A hydrodynamical theory of the initial stage of 
the vertical water impact of spheres is presented, 
which leads to decelerations throughout the impact 
stage in close accord with experiment. A dynamical 
theory of the underwater trajectory of a projectile 
in terms of a set of ten suitably chosen coefficients is 
included, together with a corresponding dynamical 
analysis of the impact phase. 


behavior of Mk 13-6 aircraft torpedo models has 
been especially carefully studied and found to agree 
reasonably well with the behavior of the prototype 
dummy. The trajectories of 1-in. to 8-in. models of 
this torpedo have been obtained at entry angles in 
the range 12 to 35 degrees. 

As an introduction to the detailed study presented 
in later chapters, an explanation of qualitative 
nature is given in Chapter 2 of the Section IV final 
report, by exhibiting a series of photographs of 
water entry which show motion of the water surface, 
the velocity of water particles throughout the 
liquid, and the form of the air cavity produced by 
the entry. The various stages of entry are clearly 


STUDIES OF WATER ENTRY BALLISTICS 


11 


seen. First comes the impact stage, of short dura- 
tion and high local pressure, during which period 
the water motion is set up. During this stage the 
water adheres to the nose of the specimen, arising 
shortly to form a thin splash sheath, as shown in 
Figure 6. Before the nose has penetrated far, 
usually less than one-half diameter, separation 
occurs between the solid and fluid, and a re-entrant 
cavity results, as in Figure 7. In the next stage the 
cavity becomes well developed about the nose and 


Chapter 3 of the Section IV final report presents 
a series of trajectories of both high-drag and low- 
drag projectiles, and gives a discussion of the 
change in form to produce desired projectile be- 
havior. Tests on antiricochet characteristics are 
included. 

In Chapter 4 a general discussion of the problem 
of modeling is given. 

Chapter 5 discusses nearly a dozen secondary 
effects which might influence model behavior. These 



Figure 7. Subsequent development of narrow entry cavity. 


may persist with a well-defined separation point, as 
in the case of bluff, high-drag nose shapes, or it may 
tend to conform closely to the nose contours, as in 
the case of fine streamlined noses. In the final stages 
of motion, the cavity seals from the atmosphere, as 
in Figure 8, and gradually closes about the speci- 
men, to be dispersed into a series of bubbles. The 
photographs presented in the report proceed from 
vertical entry of simple shapes to oblique entry of 
model projectiles.® 

e For a much larger number of photographs of similar char- 
acter, see reference 2. 


effects were investigated briefly, for the purpose of 
determining their practical significance in modeling 
high-speed water entry. They include tank-wall 
effects, adhesion, tensional stress, surface condition 
of specimen, surface tension, externally impressed 
pressure and vapor pressure, compressibility of 
fluid, change in compressibility of the solid, gravita- 
tional acceleration and flexure of specimen. Among 
other things, they show the importance of the vis- 
cosity of the air in producing under-pressure under 
the nose of the model. 

Chapter 6 presents theoretical and experimental 


12 


ANTISUBMARINE WEAPONS AND UNDERWATER BALLISTICS 


investigations of nose under-pressure with many 
details on the action of the venting. 

Chapter 7 presents a theoretical and mathematical 
treatment of the underwater trajectory of projectiles. 

Chapter 8 presents a mathematical treatment of 
the impact of a sphere on water. 

Chapter 9 gives experimental studies of the 
impact stage and includes the impact drag on 


Chapter 11 consists of recommendations for a 
continuation of investigations of this type. It deals 
with experimental and theoretical procedures and 
the development of experimental facilities. 

Appendix 1 describes the Morris Dam facilities 
in considerable detail. 

Appendix 2 describes the model and laboratory 
facilities used. 



Figure 8. Surface closure of cavity. 


spheres and cones, the tangential force on a sphere, 
the impact lift coefficient of a sphere, and the 
variations of entry whip with nose curvature. 
Observations of entry whip are by means of an 
optical lever system which gives high accuracy. 

Chapter 10 presents a large amount of experi- 
mental results on model similitude for a variety of 
specimens of widely different sizes and with special 
emphasis upon torpedo models. 


The bibliography lists all Section IV reports (117) 
classified by subject. 1 

Throughout the book are references to the char- 
acteristics of Service types of underwater ordnance, 
in the perspective of the phenomena being discussed. 


f All these reports are included in the general bibliography 
appended to this volume. There they are listed by OEMsr-418 
identification numbers, rather than by subject. 



Chapter 2 

AIRCRAFT TORPEDO DEVELOPMENT AND TESTING 


By F. C. LindvaW 


21 INTRODUCTION 

C hapters 2, 3, and 4 present in summary form 
the activities and results of the torpedo launch- 
ing group (Section VII) which operated at the 
California Institute of Technology under Contract 
OEMsr-418. Although this Division 3 contract was 
concerned primarily with rocket developments, the 
inclusion in it of torpedo studies was advantageous. 
The immediate object of these studies was the 
measurement, in full-scale launching experiments, 
of the phenomena associated with the entry of a 
torpedo into water after release from a fast airplane 
at a relatively high altitude. This work, like the 
broader Division 6 torpedo program of which it was 
really a part, had as its ultimate objectives the pro- 
viding of torpedo-plane pilots with more effective 
torpedoes and more freedom as to altitude and speed 
of flight at the time of release. 

Out of the CIT studies came the shroud ring 
modification of the Mk 13 torpedo, which demon- 
strated in combat and in tests its superior per- 
formance under the most extreme conditions likely 
to be imposed by use from present types of carrier- 
based aircraft. Other results included substantial 
contributions to the design of the Mk 25 torpedo 
and to the general art of torpedo development. 
Starting from scratch in 1943, the program involved 
development and operation of launching facilities, 
of associated photographic and other equipment for 
recording the external phenomena of entry and 
underwater run, and of torpedo-borne instruments 
for internal measurements of stresses, accelerations, 
orientation, etc. Studies of torpedo control com- 
ponents and engineering design and structural 
analysis of torpedo bodies and components were 
also included. 

Only brief descriptions of the work and its results 
are given in this summary. All aspects are covered 
completely in the final report b of Section VII. 1 

a Supervisor of Section VII (Torpedo Launching) of Con 
tract OEMsr-418 at the California Institute of Technology. 

b All earlier reports of Section VII are included in the bibli- 
ographies of the final report and of this Division 3 Summary 
Technical Report. 


22 NEED FOR IMPROVED AIRCRAFT 
TORPEDOES 

The U. S. Navy began World War II with an air- 
craft torpedo designated Mk 13. During the period 
between wars only limited experimental facilities 
were available to the Navy Torpedo Development 
Group, and little experience had been accumulated 
with this weapon. As a result, very conservative 
tactical limitations on altitude and speed of re- 
lease had been set which were serious handicaps in 
combat use. Even with these limitations the water 
entry behavior of this torpedo and the subsequent 
runs were considered unsatisfactory. Early combat 
experience with the Mk 13 in aircraft drops was 
reported as discouraging. Hooking and broaching 
occurred with distressing frequency. There was 
obvious need for aircraft torpedoes which could be 
released at higher altitudes and higher airplane 
speeds with better entry and run performance. 
Such improved torpedoes were needed not merely 
for the rather slow torpedo planes then in use, but 
even more for effective exploitation of the potential- 
ities of the faster aircraft then under development. 

Investigations toward this end were initiated in 
Section C4 of the National Defense Research Com- 
mittee. In the winter of 1942 to 1943, at the request 
of the Bureau of Ordnance, NDRC embarked on a 
two-pronged attack on the problem. First, the 
existing Mk 13 was studied with a view toward 
improvements in components and design by which 
the effectiveness of the weapon could be increased 
immediately. Second, a completely new design 
(now designated Mk 25) was undertaken. Within 
NDRC, general responsibility for this program was 
assigned to Division 6. c To Division 3 were 
assigned the essential fundamental studies of the 
hydromechanical phenomena associated with the 
entry of torpedoes into water at high speeds. All 
Division 3 work in this field was carried out by the 
California Institute of Technology, which estab- 

c For a broader account of World War II developments in 
aircraft torpedoes, see Volume 21 of the Division 6 Summary 
Technical Report. 


13 


14 


AIRCRAFT TORPEDO DEVELOPMENT AND TESTING 


lished a new section (VII) for it under Contract 
OEMsr-418. 


2 3 THE SHROUD RING TAIL 

One of the most spectacular results of the CIT 
work was the shroud ring modification of the tail 
of the Mk 13 torpedo, developed in 1944. 

During some of the early launchings of dummy 
Mk 13 units with special braces in the tail structure, 
it was observed that their additional drag greatly 
stabilized the entry, minimizing the tendency of the 
projectile to hook and to broach. Early work by 
Section IV at Morris Dam on “Mousetrap” anti- 
submarine rockets led to the use of a ring type of 
tail for stabilizing the underwater trajectory of that 
weapon. Model studies d in the CIT high-speed 
water tunnel under a Division 6 contract on a ring 
type of tail for improving the underwater stability 



Figure 1. Shroud ring Mk 1 Mod 0 assembled 
on a Mk 13-2A torpedo. 

of the Mk 13 torpedo led to a shroud ring having low 
drag in the steady run. Working with the water 
tunnel group, Section VII made and tested full- 
scale designs of such ring tails on Mk 13 torpedoes. 
One of these is shown in Figure 1. It is fabricated 
of steel with a streamlined cross section of JHrin. 
maximum thickness. The ring stiffens the guide 
vanes materially and affords protection to the tail 
structure at entry. Tests at the launching range 
demonstrated that these ring tails enabled Mk 13 

d Summarized in the Division 6 Summary Technical Report. 


torpedoes to enter the water at higher speeds, with 
less hooking and broaching, and with more stability 
in their underwater runs. 

The California Institute converted a number of 
Mk 13 torpedoes to shroud ring construction for use 
by the Naval Air Station, San Diego, in training 
drops. The response of the Ordnance Officers and 
Torpedo Squadron personnel to the improved per- 
formance was enthusiastic. Demonstration exer- 
cises were conducted against maneuvering target 
ships utilizing both standard and ring tail torpedoes. 
In compliance with a request initiated by the Air 
Station, CIT made and installed a substantial num- 
ber of shroud rings on torpedoes for issue to the 
Fleet. Subsequent tests at Pearl Harbor led to a 
request through the Bureau of Ordnance and 
NDRC for expansion of the conversion program, 
with the result that CIT modified approximately 
1,000 Mk 13 torpedoes for service use. Many of 
these went immediately into combat use. Naval 
stations, with the aid of plans and specifications 
supplied by CIT, continued the conversion program 
on a larger scale. 

Meanwhile the San Diego Naval Air Station was 
adding to the evidence of superior shroud ring per- 
formance at various launching speeds from 130 to 
300 knots and entry angles from 20 to 35 degrees. A 
report circulated in July 1944, on launchings at 
San Diego, indicated comparative performance as 
follows: 


Torpedoes dropped 
Hot and straight 
Hot and straight, with 
hooks under 25 yd 
Hot and straight, with 
hooks over 25 yd 


With Shroud Ring 
218 

204 93.5% 

199 91.2% 

5 2.3% 


Without 

358 

289 80.7% 
210 58.7% 
79 22% 


The first combat action was on August 4, 1944. 
Continuing combat experience, some involving ac- 
tions in which both standard and ring tail torpedoes 
were used, confirmed the superior performance of 
the latter, in increased percentages of hot, straight, 
and normal runs, when released from airplanes at 
higher altitudes and speeds. The ring tail modifica- 
tion of the Mk 13 torpedo was established as a 
weapon capable of withstanding any entry condi- 
tions which could reasonably be imposed by existing 
carrier-based aircraft . 

A natural consequence of the stabilization of 
water entry is that the ring tail torpedo tends to dive 
somewhat more deeply than the original Mk 13, 
which is likely to make a shallow dive followed by a 


COOPERATIVE TESTS 


15 


severe broach. As a compensating advantage, the 
ring tail torpedo can be made to enter at a flatter 
angle without damage or serious broach and thus 
achieve a shallow dive . Many tests made by the Air 
Station at San Diego in shallow water indicated 
that the deeper dive of the ring tail torpedo need 
be no tactical handicap. Studies at the Newport 
Naval Torpedo Station showed that the ring could 
be moved forward on the guide vanes to effect a 
compromise between the greater stability of the aft 
position and the instability of the bare tail structure . 
The depth of dive with the ring in the forward 
position is somewhat reduced; however, the combat 
need for this torpedo was so great that the conver- 
sion program for the ring in the aft position, which 
was already under way, was allowed to proceed so 
as not to incur the delay which further testing of 
ring position would have required. In any event 
the big improvement of the Mk 13 performance 
came from the introduction of the ring, and changes 
in performance resulting from differences in ring 
position would necessarily be small. 

24 OTHER IMPROVEMENTS OF THE 
MK 13 TORPEDO 

Further studies on the Mk 13 were directed toward 
improvement of existing components. A good deal 
of study was given to structural features, heat 
treatment of propeller shafts, studies of bearings, 
and heat treatment of propeller blades. It was 
found that the tendency for blade bending at water 
entry could be reduced by proper heat treatment of 
the existing propellers. A good deal of study given 
to the problem of gyro damage resulted in a type of 
bearing which at the close of the OEMsr-418 work 
in late 1945 showed promise of withstanding 350- 
knot entry speeds. Attention was given to the 
control with the object of eliminating some of the 
underwater roll and malfunction of the control sys- 
tem in the early stages of the underwater trajectory. 
Much of this work closely paralleled the program of 
testing components for the Mk 25. 

2 5 CONTRIBUTIONS TO THE DESIGN 
OF THE MK 25 TORPEDO 6 

The Mk 25 torpedo, which was the responsibility 
of the Columbia University group under Division 6, 

e For broader coverage of this project, see the Division 6 
Summary Technical Report, especially Volume 21. 


NDRC, represented a completely new design of 
torpedo. The launching facilities and engineering 
experience of the California Institute of Technology 
torpedo launching group were utilized to a consider- 
able degree on the structural aspects of the problem. 
Various torpedo shells were tested at the launching 
range for damage at entry, and as weaknesses ap- 
peared design changes were made. A considerable 
amount of work was done on afterbody and vane 
construction because of the new problems created 
by the use of hollow guide vanes for torpedo engine 
exhaust. The new type of joint ring evolved for the 
Mk 25 also required a good deal of structural study. 
New propellers which were designed for this unit 
were also the subject of a good many launchings. 
Cast afterbodies of various types were tested and 
commercial facilities for casting experimental alu- 
minum afterbodies were made available in the 
Southern California area to supplement the work 
which was being done in the East. As the develop- 
ment work proceeded, these additional torpedo 
components were sent to Morris Dam for launching 
tests, with particular attention being paid to the 
ruggedness of control elements of the Mk 25. 

26 COOPERATIVE TESTS 

Cooperative tests were made also for several other 
agencies. For the Applied Physics Laboratory of 
the University of Washington launching tests were 
run from time to time on a number of exercise heads 
incorporating the exploder mechanism being de- 
veloped by that group. The Allegany Ballistics 
Laboratory requested launching tests of special 
propellants to discover if the shock of entry caused 
structural damage. Special torpedo engine igniters 
were tested for the Naval Air Station at San Diego 
and for the Columbia University group. The West- 
inghouse Electric Company submitted models of an 
electric aircraft torpedo for water entry damage 
studies. This work involved not only the torpedo 
structure itself, but also detailed studies of damage 
to propellers, control gear, motor, and battery. 
Some studies of the AAF hydrobomb were made. 
For the Navy, water entry tests of the Mk 1 drag 
ring were made with and without the streamlined 
nose cap which was then under study. Also, as a 
part of the basic research study with the Applied 
Mathematics Panel, launchings were made of cer- 
tain special head shapes. 


Chapter 3 

BASIC RESEARCH ON TORPEDO ENTRANCE PHENOMENA 

By F . C. Lindvall 


A research program directed toward more basic 
information associated with the phenomena 
of the entry of torpedoes into water was carried on 
concurrently with the various aspects of the devel- 
opment work. The water entry and behavior 
studies were extensive in both theoretical and 
experimental aspects because of the large number 
of parameters involved. The studies of water entry 
were broken into five definite stages involving 
various phenomena: shock stage, establishment of 
flow, cavity stage, transition stage, and complete 
immersion. The shock stage involves the water 
forces which are the result of an acoustic shock ex- 
perienced by the body at water contact. These 
forces are extremely high and, because of applica- 
tion at an oblique angle, involve longitudinal mo- 
mentum transfer as well as angular momentum 
transfer. These forces are of extremely short time 
duration, as shown both by theoretical considera- 
tions and experimental evidence. The rotating disk 
camera b gives distance-time data which are quite 
precise. The maximum impulsive velocity change 
which could occur within the limits of error of 
measurement with this camera are of the order of 
0.5 per cent or, for typical launchings, 2 to 4 fps. 
From nose-mounted accelerometers and pressure 
plug data the magnitude of the initial shock can be 
determined, leading to time estimate for the dura- 
tion of the acoustic shock of the order 10~ 4 second. 
Transverse velocity changes due to this impulsive 
force have also been determined to be of the order 
of 2}/2 fps. However, none of these measurements 
can be considered wholly satisfactory because the 
torpedo itself is an elastic body capable of vibration 
in longitudinal and transverse modes with periods 
comparable to the time intervals under considera- 
tion. However, the evidence is good enough to 
indicate that to a considerable degree the whip at 
entry is caused by the forces during the shock stage. 
Also during this shock stage, as indicated by the 

a For another discussion not limited to torpedoes, see Sec- 
tion 1.3 of this volume. 

b This item, and other instrumentation, is described in 
Section 4.2. 


pressure plugs, on portions of the torpedo shell very 
high hydrostatic pressures exist which may cause 
local damage. To a considerable extent the Mk 1 
drag ring tends to cushion this entry shock and 
minimize local damage. The shock subjects the 
torpedo components to high acceleration forces, but 
little damage results because the various com- 
ponents are sufficiently elastic to be self-protecting 
against forces of such short time duration. 



ENTRY VELOCITY IN FEET PER SECOND 


Figure 1. Average entry deceleration as a func- 
tion of entry velocity for Mk 13 head shape (Head 
F). 

The establishment of flow is subject to a good 
deal of uncertainty because of the difficulty of ob- 
taining satisfactory detailed information during the 
first foot or two of torpedo travel into the water. 
The rotating disk camera gives deceleration infor- 
mation which is valid immediately after the very 
short time occupied by the shock stage. Typical 


16 


BASIC RESEARCH ON TORPEDO ENTRANCE PHENOMENA 


17 


data are given in Figure 1 showing a very close ad- 
herence to a square-law drag force beginning with 
the moment of head contact . This deceleration may 
be expressed as a “drag coefficient.’ ’ Figure 2 shows 
the variation of the drag deceleration with time 
after entry, assuming constant drag coefficient. 
This drag coefficient is substantially constant for 
full torpedo immersion and one or two lengths of 



Figure 2. Drag deceleration vs time after entry. 


additional travel. Then, because of effects occurring 
in the cavity stage, a higher value of drag coefficient 
is observed followed by the normal low body drag 
appropriate to fully immersed travel at speeds below 
the cavitation velocity. Figure 3 shows this effect. 

As seen from Figure 2, the high values of accelera- 
tion may exist for 0.1 second or more. Internal 
components of the torpedo are thus subjected to 
high forces which last for periods of time which are 
large compared with their own natural periods . As 
a result, for all practical purposes, all but very 
flexibly mounted components are subjected to static 
loads corresponding to these high accelerations. 

In the cavity stage of entry the torpedo is thought 
to be in unstable balance on its nose with the tail 
structure moving transversely in some direction 
through angular momentum acquired in the initial 
stage of entry. Sooner or later the tail structure 


encounters the more or less solid water which 
bounds the cavity, with resultant tail slap and 
application of hydrodynamic forces. The shape of 
the surfaces on the tail structure may cause the tail 
to dig into the wall of the cavity. The exact nature 
of this behavior is not known for full-scale tor- 
pedoes, but model studies (see Section 1.3) have 
indicated the performance as described to be typical. 
In full-scale tests the acoustic range records show 
evidence of this tail slap occurring well after the tail 
has disappeared below the surface of the water. 



0 50 100 150 

DISTANCE IN FEET 


Figure 3. Mean velocity-distance curve for 
dummy aircraft torpedoes. 

Pressure plug data also indicate high values of im- 
pact pressure on portions of the tail structure and 
the afterbody. A considerable amount of damage 
due to this tail slap has been observed in afterbody 
shells. 

The transition stage from the cavity state to that 
of complete immersion or wetting of the torpedo 
can only be inferred for the full-scale torpedoes. 
The model work (see Section 1 . 3) shows the cavity 
to be followed by a bubble which breaks up until 
finally the torpedo is fully wetted. The acoustic 
range gives some evidence of sounds which are inter- 
preted as bubble collapse, and general photography 
shows the position at which entrained air finally 
reaches the water surface. The observed position of 
rise of the bubbles correlates well with the measured 
information on drag coefficient change from high to 
low value. 

In the complete immersion stage the underwater 
trajectories were carefully determined by acoustic 
range data and actual perforations of nets along 
trajectory. These data, together with the known 


18 


BASIC RESEARCH ON TORPEDO ENTRANCE PHENOMENA 


positions of entry and broach, if any, gave very 
satisfactory trajectory records of the type of Figure 
4. These underwater trajectories were investigated 
for effect of velocity, pitch, yaw, and roll of the 
torpedo at entry. The data in Figure 4 show the 
general trend of the trajectories as affected by entry 
velocity. The general effect of the initial roll was 
slight except for conditions of large amounts of 


including a group of sphere-ogive combinations pro- 
posed by the California Institute of Technology 
Hydrodynamics Laboratory, none gave a signif- 
icantly better performance than the Mk 13 head in 
resisting a dive to the bottom due to steep pitch at 
entry. Included in these head studies was one con- 
sisting of the Mk 13 shape to which was added a 
90-degree cone. This cone, shown in Figure 12 of 


UJ 

UJ 


z 

X 

H 

CL 

UJ 

O 


HORIZONTAL DISTANCE IN FEET 

0 20 40 60 80 100 120 140 160 180 200 


















1 

400 

1 1 

FPS 












500 

FPS 




330 

FPS 


































































Figure 4. Underwater trajectories of Mk 13 dummy aircraft torpedo for initial pitch between 1 degree steep 
and 1 degree flat. Numbers of launchings: 330 fps, 16; 400 fps, 3; 500 fps, 7. 


rudder setting. The effect of yaw was much the 
same as that of pitch, except of course in inducing 
horizontal deviations from a straight-line trajectory. 

Pitch, that is the angle made by the longitudinal 
axis of the torpedo with respect to the trajectory, 
had a marked effect on the depth of dive or the 
tendency to broach. Figure 5 is typical of many 
sets of data taken for the purpose of showing the 
sensitivity of a particular head shape to the amount 
of pitch at entry. The data are shown in two ways: 
in the upper curve, the deviation of the trajectory 
from a straight-line projection of the air flight tra- 
jectory is measured at an arbitrary distance of 100 ft 
from point of entry. The lower curve gives the 
absolute depth of dive as a function of the pitch at 
entry. For the particular head shape used in these 
tests a steep pitch of 2 degrees or more leads to deep 
dives, and as much as 3 degrees of steep pitch would 
put the torpedo on the bottom except in very deep 
water. Flat pitch on the other hand leads to shallow 
dives, but no abnormal behavior, in the sense of an 
excessive broaching tendency, is indicated. Another 
presentation of data of this type is given in Figure 6, 
in which the actual trajectories are given with ap- 
propriate legends indicating the number of degrees 
of flat or steep pitch and the number of launchings 
of nearly the same amount of pitch which have been 
grouped as a single composite trajectory. 

Although a variety of head shapes was tested, 



DEVIATION AT 100 FT FROM LINEAR PROJECTION 
OF AIR FLIGHT TRAJECTORY AS A FUNCTION OF PITCH 



Figure 5. Pitch sensitivity of Head K. CIT full- 
scale dummy aircraft torpedo. 


BASIC RESEARCH ON TORPEDO ENTRANCE PHENOMENA 


19 


Chapter 4, does not improve the tendency to dive, 
but increases the broaching tendency for flat pitch 
and definitely introduces a larger whip, which is 
undesirable from the standpoint of structural dam- 
age. No very large departure from Mk 13 dimen- 
sions was made in any of these heads because of the 
overall torpedo length, which was fixed by aircraft 
limitations, and the necessity for maintaining ap- 


the total weight was maintained constant and 
moment of inertia held fixed. The center of gravity 
positions were fore and aft with respect to the center 
of buoyancy and transversely, above and below, 
with respect to the longitudinal axis of the body. 
With the center of gravity forward of the center of 
buoyancy, greater entry stability was demonstrated 
although the underwater trajectories and depth of 


HORIZONTAL DISTANCE IN FEET 



Figure 6. Underwater trajectory as a function of pitch for Head K. CIT full-scale dummy aircraft torpedo. 


proximately the same war-head volume . The varia- 
tions in shape were more significant with respect to 
steady running drag and cavitation parameter than 
in modifying the pitch sensitivity. 

With various dummies, some of which were also 
used in the establishment of the underwater trajec- 
tories, the entry and underwater performances were 
investigated with respect to shroud ring size and 
reaction, rudder setting, length-to-diameter ratio, 
trim, and moment of inertia. The most extensive 
work related to the trim and moment of inertia 
studies. In the trim studies the center of gravity of 
the body was adjusted to different positions while 


dive tended to be greater. In the moment of inertia 
studies the weight and center of gravity position 
were held fixed while the moment of inertia about 
the center of gravity was varied. The effect of the 
moment of inertia is not striking within the limits 
that are physically possible in a torpedo, but, in 
general, the greater the moment of inertia, the less 
violent are the actions of the torpedo at entry. The 
effect of greater length-to-diameter ratio is not 
entirely independent of the moment of inertia, 
which inevitably increases, and is similar in that 
trajectories are obtained which tend to follow more 
nearly a projection of the airflight path. 


20 


BASIC RESEARCH ON TORPEDO ENTRANCE PHENOMENA 


A large part of the study of underwater trajectories 
with dummies was made for the purpose of correlat- 
ing, if possible, the observed performance with 1-in. 
models being studied by Section IV of Contract 
OEMsr-418. In these comparative studies the model 
and prototype dynamic properties were carefully 
scaled, and the velocities, model and prototype, were 
related through Froude’s rule . No attempt was made 
in this model work to vary the pressure of the at- 
mosphere above the water. Although it was found 
that the trajectories of prototype and model cor- 
related in a general way in the early stage of the 


underwater run, significant deviations were ob- 
served as model velocities became low. More sig- 
nificant, however, was the radically different be- 
havior of certain head shapes which, with the 
model, dove consistently to the bottom, while the 
prototype followed normal trajectories with upward 
curvature. Using pitch sensitivity of different heads 
as an index, the correlation between the model and 
prototype behavior was unsatisfactory. 0 

c Section 1 .3 of this volume indicates that better correla- 
tions may be obtained by modeling on a velocity basis instead 
of by Froude’s rule, and by venting the models. 


Chapter 4 


FACILITIES AND INSTRUMENTATION FOR STUDY 
OF TORPEDO ENTRY 

By F. C . Lindvall 


41 GENERAL FACILITIES 

T he first problem of the CIT torpedo group 
was the design and installation at a suitable 
site of equipment capable of launching torpedoes 
into water at velocities and entry angles correspond- 
ing to high-speed aircraft drops. Among the 
requirements were sufficient water depth and length 
of run for adequate observation of the effects of 
interest. By arrangements made earlier in connec- 
tion with other CIT underwater ordnance investiga- 
tions, the Institute had a suitable site available only 
20 miles east of Pasadena, on the artificial lake 
above the Morris Dam, owned by the Metropolitan 
Water District of Southern California. This site met 
these requirements in that it provided a 5,500-ft 
straight course of depth 100 to 140 ft. All the 
present and projected launching equipment is 
located on a peninsula approximately 3,000 ft up- 
stream from the dam. This peninsula has a steep 
slope which provides a convenient support for 
mounting a launching tube. Steep mountains near 
the torpedo entry point provide excellent locations 
for detail and general view camera stations. The 
mild climate allows work to continue throughout 
the year, with good photographic conditions on 
almost all days. 

Various schemes for accelerating and launching 
the torpedo were studied, leading to a final decision 
for the construction of a 300-ft tube for compressed 
air launching. It was believed that sufficient useful 
information could be obtained with a tube of fixed 
entry angle having the diameter of the existing tor- 
pedo to justify immediate construction of this 
facility, without incurring the considerable loss of 
time which would be required for the design of a 
more elaborate launcher to accommodate other pro- 
jectile sizes and permit adjustable angle of entry. a 
An entry angle of approximately 19 degrees was 
chosen to match the general limits proposed by the 
Bureau of Ordnance for 350-knot airplane speed and 

a A variable angle launcher of CIT design was added to the 
facilities after they were taken over by the Navy in 1945. 


800-ft altitude of release. This angle was fixed 
with the realization that the corresponding water 
entry angle would probably be the lower limit of 
tactical operation at which satisfactory entry could 
be obtained and for which also torpedo damage at 
entry would be accentuated. 

Design work on the launcher and associated facil- 
ities began early in 1943; construction of buildings 
and foundations at the site, early in the summer of 
1943, concurrent with fabrication of launcher com- 
ponents. The equipment was installed during the 
summer and the first launchings were made in 
August 1943. The launching facilities have been in 
continuous use since that time and are now being 
operated on a permanent basis by the Underwater 
Ordnance Section of the Naval Ordnance Test Sta- 
tion, Inyo kern. During this period the facilities 
underwent continuous improvement as the results 
of the research program dictated modifications and 
additions . 

The general problems set for the CIT torpedo 
launching range were as follows: 

1. General hydrodynamic effects at entry. 

2. The effect of dynamic characteristics of the 
torpedo. 

3 . The effect of nose and tail structures on entry 
and underwater trajectories. 

4. The determination of underwater trajec- 
tories . 

5. The measurement of deceleration forces and 
the effects on structure and mechanisms of the con- 
sequent impact loadings. 

6. The general structural aspects of the entry 
problem. 

Figure 1 is a view of the range from a point 
directly over the launching tube. The two lines of 
buoys in the foreground are 100 ft apart and serve 
to support an array of hydrophones which con- 
stitute the acoustic range. In the distance may be 
seen a set of six sonobuoys which serve to extend 
the acoustic range for tracking the torpedo on its 
run. At the left in the foreground are located a 
control station and a camera car which is positioned 


21 


22 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 


opposite the point of water entry and is moved 
along an inclined track parallel to the launching 
tube to follow changes in water level. Figure 2 is a 
sketch map of the facilities on the peninsula. Figure 
3 is a plan and elevation of the launching tube itself 
in relation to the torpedo shop and working areas. 


thereby giving an immediate unrestricted flow of air 
from the tank into the tube aft of the torpedo imme- 
diately the torpedo had moved forward sufficiently 
to clear the inlet from the tank. 

Figure 5 is a schematic drawing of the tube 
launching mechanism. In operation, after the 



Figure 1. General view of launching range area looking down range. 


Figure 4 shows the breech end of the tube and the 
Y-connection to the compressed air impulse tank. 
A large pressure vessel of 1,100 cu ft capacity and 
rated working pressure of 150 psi, which was on the 
California Institute campus, was made available for 
this project to permit speedy completion of launch- 
ing facilities. Later, as materials and fabrication 
facilities for pressure vessels became obtainable, a 
new tank of greater capacity and higher pressure 
rating was built to replace this item of Institute 
equipment. After a study of possible quick-acting 
valves for release of impulse air, a decision was 
made to have the torpedo act as its own plug valve, 


breech door has been closed and ring seals (8) have 
been pressurized, the main gate valve (9) is opened. 
Any leakage into the breech section is vented 
through valve (5) until launching is desired. The 
torpedo is held in position by detent pin (7) . When 
release is desired, handle (6) is pulled, which simul- 
taneously releases the detent pin and seal pressure 
and closes vent valve (5) . Air passing by the after 
seal (8) pressurizes the breech section, causing the 
torpedo to move forward in the tube and clear the 
Y-connection for full release of impulse air. The 
breech door is shown in detail in Figure 6. 

The performance of the launching tube is shown 



GENERAL FACILITIES 


23 


in Figure 7, which gives calculated and measured 
velocities — two for the standard Mk 13 torpedo and 
one for a 1,500-lb dummy — at various distances 
along the 300-ft launching tube for three values of 
impulse tank pressure. Subsequently at two sta- 
tions along the tube were added a series of gas 
booster tubes. These utilized standard rocket 


to the trajectory) an air scoop was added at the 
muzzle of the launcher which could be swung out of 
use or changed from top to bottom location. 

The Mk 13 aircraft torpedo served as a utility 
instrument in many of the launching tests of this 
project, as well as being an object of study for 
possible improvement for immediate service ap- 



® CAMERA CONTROL STATION 
(2) FLARE CAMERA 

® SIDE VIEW MOTION PICTURE CAMERA _ 

(4) REAR «» it « H cHg) 

® GENERAL ^IDE VIEW M P C 

© » REAR » 'I 

® » VIEW 


1 GUARD HOUSE 
WATER. TANIA 
TOILET BLDG. 
WATER, SUPPLY PUMP 
OFFICE BLDG 
MACHINE SHOP 
BREECH 
UTILITY BARGE 
STORAGE 


MAGAZfNE 

CARPENTER’S SHOP 
LOADING RAMP 
BOAT HOUSE 
SOUND RANGE 

SOUND RANGE RECORDING BUILDING 


O 40 80 
U_l l 

SCALE IN FEET 


Figure 2. Plan of the launching site showing principal facilities. The elevation of the highway is 1,321 ft, 
the elevation of the working area (E, F, and G) is 1,305 ft, and the elevation of the water surface is nor- 
mally 1,160 to 1,167 ft. 


motors as sources of additional high-pressure gas, 
injected just as the torpedo passed each station. 
The effect of these boosters was to add some 50 fps 
to the muzzle velocity of the torpedo. The need for 
the rocket boosters disappeared with the installa- 
tion of the large impulse tank, which has a volume 
of 1,550 cu ft and a working pressure of 350 psi. 

The launching tube met the general design speci- 
fications for entry angle and velocity and put the 
torpedo into the water with very small amounts of 
random pitch and yaw. Later, to induce three or 
four degrees of up or down pitch at entry (relative 


plication. Dummy Mk 13 units were also obtained 
from the Navy for some of the early launching 
trials, although these dummies were not usable for 
elaborate tests because of the impracticability of 
installing instruments. A number of dummy units 
were constructed for the project with interchange- 
able heads, center sections, and afterbodies so that 
instruments could be installed, shape modifications 
made, and dynamic properties of the body changed. 
These dummies also were made structurally rugged 
enough to withstand the maximum velocities of 
entry anticipated for this work. Dummies have 



24 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 


been launched with entry velocities up to 800 fps. 

Torpedoes are recovered in the buoyant state by 
boat and are towed to a landing ramp where they 
are floated onto a submerged trailer, which is then 
pulled ashore and on up to the torpedo shop. This 
procedure is not only rapid but is also flexible 
enough to follow the changes in lake elevation. A 
number of launchings are made with torpedoes or 
dummies in the buoyant condition. Other units are 
launched with water ballast and blowing means fol- 


operations, a battery of air compressors for launch- 
ing and torpedo-charging air, a small instrument 
and gyro laboratory, dark rooms for photographic 
work, a wood shop for construction of miscellaneous 
test equipment, a small magazine for storage of 
miscellaneous explosive material, and a limited 
amount of office space for the range supervisory 
personnel. In addition, a structure for housing the 
electronic equipment associated with the acoustic 
range is located near the point of torpedo entry. 



Figure 3. General plan and longitudinal section view of the Morris Dam Hydrodynamics Station launching 
equipment. 


owing general Navy torpedo exercise practice. For 
greater flexibility in the use of water ballast, high- 
density liquids are sometimes employed, the most 
satisfactory being a Bentonite suspension as used in 
the preparation of high-density mud for oilwell 
drilling. 

Torpedoes which failed to float after launching 
were recovered by the 11th Naval District Mine 
Disposal Unit with magnetic location and diving 
operations. The nature of the lake bottom required 
a precise location before the diver was sent down. 

Among the miscellaneous service facilities are a 
torpedo shop for overhaul and minor mechanical 


The acoustic range consists of an array of twelve 
hydrophones, as shown in the sketch of Figure 8. 
These hydrophones respond to sound impulses 
generated in one of the hand holes of the torpedo 
by detonation of electric primers set off sequentially 
by a timer. The responses of the twelve hydro- 
phones to these sounds are amplified and recorded 
simultaneously with a twelve-channel oscillograph 
which superimposes timing lines on the record. 
From the difference in time of arrival of the sound 
at the different hydrophones, the position of the 
torpedo at the moment each sound is produced can 
be computed. The reduction of the acoustic data is 


INSTRUMENTATION 


25 


made on a mechanical computer which not only 
minimizes the labor of computation , but also makes 
the best average from the redundant data. Figure 9 
gives a typical underwater trajectory in plan and 
elevation as determined from the acoustic range 
and from nets located in the range. From such a 


urements involves many difficulties of application 
and of final interpretation of results; consequently, 
except for local effects within the torpedo due to 
impact or shock loading, an external measurement 
of the behavior of the body as a whole is most satis- 
factory. During the entry phase external photog- 



Figure 4. General view of breech end of launching equipment. This illustration shows the No. 1 impulse 
tank and the breech and Y sections prior to reinforcing. 


record, correlated with the sound of water impact, 
distance-time information is obtained from which 
velocity and underwater deceleration may be 
derived . 

42 INSTRUMENTATION 

External Observations 

Measurement of the deceleration of the torpedo 
as it enters the water is of fundamental importance. 
Internal recording equipment for deceleration meas- 


raphy can follow the entry for a distance representing 
roughly 80 per cent of the body length. Ordinary 
motion picture camera technique gives inadequate 
time resolution and ultra-rapid cameras are inap- 
plicable because of insufficient light. A new tech- 
nique therefore was developed for this work, con- 
sisting of the use of a small but intense light source 
mounted on the rear structure of the torpedo, the 
image of which was recorded on a photographic 
plate every thousandth of a second by the chopping 
action of an interrupting disk in the light path of the 
camera. This disk had narrow radial slits carefully 


26 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 


spaced around the circumference which permitted 
the camera to see the light source on the torpedo a 
thousand times each second for intervals of approx- 
imately 20 microseconds each. A typical record 
obtained with this camera is shown in Figure 10, in 
which two light sources were employed on the tail 
of the torpedo. The general illumination of the 


of torpedo release by means of an electric primer 
and a bit of black powder paste. This arrangement 
gave a brilliantly illuminated slit approximately 
y 8 in. wide and 1J^ in. high. The original 5-by-7 
glass photographic plates, from one of which Figure 
10 was reproduced, were measured with great pre- 
cision on a measuring engine such as is used with 


X) bP-TECH DOOP-, 

2) LAUNCHING TUbE 

3) ^TAIVTING LANYARD LEVEE- 

4) HANDHOLE 

5 ) VENT VALVE 

g> ^ T A I7.T I N (x HANDLE 
7) DETENT PIN 
® ^EAL 

g) GATE VALVE 

g> 1550 CU FT IMPULSE TANIA 

© $H70UD E.ING DETENT PIN 



LAUNCHING 
TUBE PROPER 


TUbE LAUNCHING 
MECHANISM 

<C ALE IN INCHED ■ ■ mm 

oi 4 8 n l& 


Figure 5. Sectional view of breech end of launching equipment. Structures connected with openings 4 and 
5 are actually located 90 degrees toward reader from position shown. 


background is sufficient to bring out reference 
marks which aid in the reduction of data, but the 
total time during which the camera shutter is open 
must be kept to a minimum to avoid overexposure 
of the background. In so far as the essential record 
of torpedo position as a function of time is con- 
cerned, the images of the light source on the torpedo 
are sufficient and could be obtained at night just as 
well. The light sources used consisted of small steel 
cups with suitable mounting brackets. These cups 
were slotted and packed with an aluminum powder 
pyrotechnic mixture which was ignited at the time 


spectrograms. The camera was located approxi- 
mately 70 ft from the point of entry, and the 
precision of measurement was such that the position 
of a good flare image could be determined at the 
torpedo to within a tenth of an inch. 

By measuring the intervals between flare images 
and plotting these measurements against time, a 
velocity-time curve (Figure 11, upper curve) is ob- 
tained for the entry of the torpedo up to the time 
the tail disappears from view. These velocity curves 
form a straight line parallel or nearly parallel to the 
time axis until the torpedo strikes the water; at this 




INSTRUMENTATION 


27 


point the line joining the points bends abruptly 
downward. The slope of this portion of the velocity 
curve is proportional to the deceleration of the 
torpedo. When two flares are used on the torpedo, 
one above the other, a line joining the upper and 
lower flare images is a measure of the angle of the 
axis of the torpedo. Thus both the pitch angle and 


which permits a photographic check of the velocity 
of entry to be correlated with the muzzle velocity as 
measured with an electronic timer. 

Directly under the breech of the launching tube a 
16-mm, 64-frame-per-second camera is located to 
record the rear view of the entry phenomena. From 
these records may be determined the roll and yaw. 



Figure 6. View of breech end of launching tube. 


change in pitch angle of the torpedo can be deter- 
mined . 

Other photographs of water entry were made with 
motion picture cameras located in various positions. 
A Mitchell 35-mm high-speed (125 frames per 
second) motion picture camera was located adjacent 
to the rotating disk camera and gave in considerable 
detail from this side view the aspect of the torpedo 
prior to entry and the behavior during entry, as 
shown in Figure 12. The synchronously rotating 
timing disk in the foreground gives a time scale 


A general view motion picture camera is situated 
on the hillside approximately 400 ft from the center- 
line of the range. The field of view is about 400 ft 
at the center of the range. This camera is used to 
record general data such as positions of entry, 
broach, and re-entry, the velocity at broach, length 
of underwater run and of air travel, the angle at 
broach and re-entry, and of hook at broach, the 
height of the broach, and the path of run following 
re-entry. A timer in the field of view of this camera 
has four disks driven by a synchronous motor at 


28 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 


speeds of 1,500, 150, 15, and 1.5 rpm. Film from 
this camera is viewed for measurement with a single- 
frame projector and a system of plane mirrors which 
puts the image on a measuring grid ruled in perspec- 
tive to represent the true coordinates of the lake 
surface and aligned by placing the images of the 





DISTANCE ALONG LAUNCHING TUBE IN FEET 


Figure 7. Calculated and measured projectile 
velocities along the launching tube. 


range buoys in coincidence with their respective 
positions on the grid. 

An overhead camera may be used, if desired, in a 
camera car on a cable suspension system directly 
over the range, permitting location of the camera 
directly over the point of entry. The camera car 
and camera mechanism are remotely operated from 
the central camera control station. 

Underwater photography has been used in an 



experimental way with motion picture cameras in- 
stalled in watertight submerged drums. Only at 
certain times of the year is the clarity of the water 
sufficient to permit photography of full-scale tor- 
pedoes, because of the distance the camera must be 
located from the line of the underwater trajectory 
in order to keep a field of view great enough for more 
than a single torpedo length. 

A general rear view camera is used to record 
powered runs of the torpedo. This is a single- 
exposure camera which, from a height above the 
lake surface, photographs the track of the torpedo 
at any desired stage in the run. The photographs 
are measured by placing them over a transparent 
grid so that the coordinates of the torpedo track 
may be determined. 


422 Internal Measurements 

Internal instrumentation included a variety of 
devices for obtaining acceleration of torpedo com- 
ponents, pitch, roll, propeller speed, control posi- 
tions, time of water entry, and miscellaneous events 
to be correlated with the moment of water entry. 
The heart of the recording system for these various 
instruments was a specially designed neon tube 
camera, as shown in Figure 13, using a 725 -watt 
neon bulb as the essential element. Three models 
of this recording camera have been constructed and 
used. Each unit consists essentially of a bank of 
neon bulbs, an optical system which projects the 
light onto moving motion picture film, a film drive, 
a vacuum tube oscillator which periodically flashes 
one of the neon tubes and thus produces a timing 
reference trace on the film, and switches which start 
and stop the camera. In the various models bat- 
teries are either self-contained in the camera or are 
placed in an auxiliary box in the torpedo. All this 
equipment must be extremely rugged in order to 
avoid distortion or damage resulting from the 
severe shock of high-speed water entry. The neon 
bulbs are either on or off depending on contact 
position in the instrument whose operation is being 
recorded. A typical record obtained with this cam- 
era attached to a step accelerometer is shown in 
Figure 14. 

Accelerometers 

This step type of accelerometer consists of a series 
of cantilever springs, shown schematically in Figure 





INSTRUMENTATION 


29 



Figure 8. Aerial perspective showing torpedo launching area and sound range as of July 1945. 


£nc/ observed . 


20.6° OATE 7-11-4-5 

0.2* f RUDDERS 

O.O* Horiz J up - 4 / 'down 

- 0.9 * Vert Top 3.S ^Bottom .2.6 

REMARKS Hot s hoi. Depth setting JO feet Gyroscope prespup.^/ Wingless tail. 


LAUNCHING NO J8S7 
TORPEDO MX 13-2A 
ENTRY V IN FPS 34-1 
WEIGHT IN LB 2053 


ENTRY ANGLE 
PITCH ANGLE 
YAW ANGLE 
ROLL ANGLE 


Single yaw rocJcef set for 2° nose right yew. Test/ of new electric net et Station 477. 


LEGEND 
A ENTRY & 2H4 SOUND DATA 
O CAP DATA 
Q NET DATA 

SURFACE POINTS 



Figure 9. Typical trajectory plot. 



LAUNCHING NO. IBS 7 



PITCH IN DEGREES VELOCITY IN EPS 


30 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 



Figure 10. Flare camera record with two flares on torpedo tail. 



Figure 11. Typical velocity-time (upper) and 
pitch-time (lower) curves derived from flare meas- 
urements. 


15, preloaded to break contact under a specified 
value of acceleration and each controlling one of the 
neon camera bulbs. The record of Figure 14 is 
actually for three separate accelerometers located in 



Figure 12. Entry of Mk 13 torpedo with added 90- 
degree nose cone. 


the torpedo at the three positions shown in the 
sketch. The record shows the 500-c timing marks, 
auxiliary 100-c timing marks, and the responses of 
the 21 neon tubes connected to the three accelerom- 








INSTRUMENTATION 


31 



Figure 13. Neon tube camera — Model 3. Top and bottom plates removed, showing removal of neon tube bank. 


STEP ACCELEROMETER 
SPRING SETTINGS 


START OF NOSE 
TRANSVERSE ACCELERATION 


500 CYCLE/SEC TIMING LINE 


INST NO. 0 

INST NO. 1 

INST NO. 2 

20 q 

20g 

20 g 

50 

50 

50 

100 

150 

100 

150 

225 

200 

250 

300 

300 

350 

400 

450 

. 400 

450 

500 


START OF NOSE 
,IAL ACCELER 

BOTTOMING OF CAMERA IN SHOCKMOUNT SYSTEM 


100 CYCLE/SEC TIMING LINE 

TIME FROM INITIAL IMPACT 0 
IN SEC | 

APPROXIMATE DISTANCE SCALE O 
IN FT 


START OF AFTERBODY 
TRANSVERSE ACCELERATION 


TAIL SLAP SHOCK AT TORPEDO NOSE 



OJOI 0t02 0.03 0.04 0.05 006 0X>7 0.08 a09 OHO O.ll 

1 I 1 

K) | 20 

ONE TORPEDO LENGTH 


30 

TWO TORPEDO LENGTHS 


ENTRY VELOCITY 
358 FPS 


STEP ACCELEROMETER IN 
AFTERBODY INST NO. 2 



STEP ACCELEROMETER IN 
NOSE - AXIAL INST NO. O 


STEP ACCELEROMETER IN 
NOSE- TRANSVERSE INST N0.1 


Figure 14. Typical step accelerometer record from a launching test. 




32 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 


eters. From this record can be obtained the time 
duration of the various magnitudes of acceleration 
shown by the blacked-out portions of the neon tube 
records. The approximate distance scale shown on 
the figure is derived from external photographic and 
underwater acoustic data. 

Various types of accelerometers of the indenter 
type and the copper ball deformation type have 
been used in the course of this work. However, the 
recording step accelerometer is by far the most 



Figure 15. Representative CIT accelerometer 
spring element. 


satisfactory, because it not only gives a time record, 
but also records a sequence of repeated shocks as 
contrasted with a single record resulting from a 
deformation or displacement type of instrument 
which gives no time history and which may be quite 
ambiguous as a result of repeated shocks of unknown 
character. For certain types of testing, the simpler 
instruments of the indenter or deformation type 
may be calibrated against the step accelerometer 
and then used with some confidence for subsequent 
accelerations of the type for which the calibration is 
valid. The step accelerometer lends itself to analysis 
and reliable calibration so that peak values of 
acceleration may be measured with confidence and 
repeated shocks determined reliably, provided the 
repetition rate is slow compared with the natural 
frequency of the spring elements in the accelerometer . 

Accelerometer Calibration 

To provide a calibrating system for accelerom- 
eters, a drop table with control and recording equip- 
ment was constructed. A table carrying a standard 
accelerometer, to which other apparatus or acceler- 


ometers for calibration purposes could be attached, 
was arranged on guide rails to have a free fall of 
approximately 20 ft onto buffers, dash pots, lead 
plugs, or other suitable stopping means. The stand- 
ard accelerometer consisted of a spring-mass sys- 
tem in which the spring was a thin-wall Dural 
cylinder to which was attached wire strain gauges 
determining the deflection of the spring system. 
The unit was calibrated statically and dynamically . 
The dynamic calibration was made by the sudden 
release of a known weight suspended from the bot- 
tom of the accelerometer. This procedure caused 
the same resistance change in the strain gauges as 
sudden loading, though with opposite sign. The 
natural frequency of this accelerometer was approx- 
imately 5,000 c and was valid therefore for accelera- 
tion measurements on phenomena of frequencies up 
to 1,500 c at least. The electrical output of the 
strain gauges was amplified and recorded on a mov- 
ing film oscillograph consisting essentially of a film 
drive and a cathode ray oscillograph beam swept 
in only one direction. A record of a step accelerom- 
eter calibration made with this equipment is shown 
in Figure 16. 

Damage Instruments 

Additional dynamic studies were made in the 
torpedo models with what were called “damage 
instruments.” These instruments consisted of 
simple mechanical structures, cantilever beams, and 
tension specimens, which were loaded by accelera- 
tion forces. Figure 17 illustrates the tension type. 
These were for standard tension specimens whose 
properties were known from static tests on similar 
components. They are secured at one end to mount- 
ing structure and loaded by acceleration forces act- 
ing on the weight attached at the other end. The 
weights are loosely guided in the enclosing cylinders. 
Figure 18 indicates an array of cantilever members, 
some of which are loaded with definite weights 
applied at the ends, others of which are uniformly 
loaded by the acceleration forces. As a result of a 
particular launching some of these test members 
are undamaged, others have taken permanent set, 
and for some of the tension specimens actual failure 
may have occurred. The information resulting from 
these tests is helpful in designing structural com- 
ponents of the torpedo to withstand the shocks of 
water entry. 



INSTRUMENTATION 


33 


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400 9 




1 1 1 

1-0.002 SEC 

— 1 

N0.6 TRACE 

300 9 




Hu 





NO. 5 TRACE 

250 9 









N0.4 TRACE 

1 

200 9 









N0.3 TRACE 

1 

150 9 









N0.2 TRACE 

1 

100 9 









N0.1 TRACE 

1 

50 9 










1 









PLOT OF STEP ACCELEROMETER RECORD 


Figure 16. Comparative plots of the wire strain-gauge accelerometer and step accelerometer records of the 
same acceleration pulse. 


Additional damage information was obtained 
through the use of scratch gauges of the de Forest 
type and from bonded-wire strain gauges. The use 
of the wire strain gauges was limited to tests in 
which external recording could be used, whereas the 
de Forest strain gauges were installed at many 



Figure 17. Tension-type damage instrument. 


points in the torpedo and at various points of its 
mechanism to determine the maximum strains 
resulting from launchings. 

Hydro Pressure Plugs for Localized Peak 
Pressures 

The typical accelerations for the torpedo as a 
whole determined from these various instruments 
lead to an interpretation of water drag forces pro- 
portional to the square of the velocity, but certain 
local damage effects are traceable to transient water 
pressure forces of much greater magnitude. To 
study this effect hydropressure plugs were de- 


veloped, as shown in Figure 19. These plugs, of 
3^-in. overall diameter, were inserted at various 
stations in the torpedo shell as indicated in the 
sketch of Figure 20. The recording is done by per- 
manent set of an annealed phosphor-bronze dia- 
phragm. This permanent set is correlated with 



Figure 18. Model No. 1 bend-type damage instru- 
ment. 


hydraulic pressure studies of identical disks. As 
seen in the tabulated data of Figure 20, extremely 
high hydrostatic pressures occur during entry which 
correspond quite well to the calculated pressures 
which are the product of the local torpedo velocity 
at impact and the acoustic impedance of the 
water. At various points on the torpedo body, 


34 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 



1-20 NF THREADS 


Dl APHRAGM 



-24 N F THREADS 


SECTION 



END VIEW 



Figure 19. Hydropressure plug and suggested mounting method. 




hjs'n? '*£'•£-£ 


STATION 

01 APHRAGM 

THICKNESS 

DEFORMATION 

PRESSURE 

P S 1 

A 

.005" 

.0231 

1,450 

B 

.005" 

.0329 

2,100 

C 

.015- 

.0317 

6,100 

D 

.015" 

.0911 

14.500 

E 

.0214- 

.0617 

20.400 

P 

.015- 

.0606 

10.600 

S 

.005- 

.0371 

2.350 

H 

.005" 

.0001 

125 

I 

.005" 

.0337 

2,150 

J 

.005" 

.0270 

1.725 

K 

.005" 

.0226 

1,425 

L 

.005" 

.0292 

l'875 

11 

.005" 

.0452 

2,825 

W 

.005" 

.0240 

1.525 

0 

.005" 

.0390 

2.475 

P 

.005" 

.0308 

1.975 

Q 

.005" 

.0029 

275 

R 

.005" 

.0285 

1.825 

S 

.005" 

.0383 

2 ,425 

T 

.015- 

T 03A6 

6,600 

U 

.0214" 

.0232 

6.700 

V 

.0214" 

,0673 

19,300 

w 

.015" 

.0319 

6,100 

Zp 

.005" 

.nom 

125 

z S 

.005" 

.0000 

125 OR LESS 



PLAN VIEW OF TAIL 


LAUNCHING NO. 613 
VELOCITY 545 FPS 
ANGLE OF ENTRY 20°23’ 
YAW 

PITCH +0*38* 

ROLL AT ENTRY — 
WEIGHT 1534 lbs. 


Figure 20. Typical data sheet and pressure distribution plots obtained with hydrophone units in a launch- 
ing test of a Mk 13 dummy torpedo at 20-degree entry angle and 550-fps entry velocity. 


INSTRUMENTATION 


3. r 


particularly in the area of nose contact with the 
water, the geometrical volume displacement of the 
water requires a water velocity in excess of that of 
the acoustic velocity. Consequently, the water is 
compressed, and high local pressures result. Similar 
effects, due to slap on the side of the entry cavity, 
occur in the afterbody and in portions of the tail 


Orientation Recorders 

A gyroscopic orientation recorder was developed 
to determine the orientation of the torpedo with 
respect to its trajectory. Knowledge of this orienta- 
tion is of great importance in determining the sub- 
sequent motion, beginning with the precontact 




Figure 21. Hydropressure plug data showing peak pressure distribution on torpedo nose with various cov- 
erings. Entry angle 20 degrees. Entry velocity 410 fps. 


structure. While these effects are of very short 
time duration, they frequently give rise to local shell 
damage. Figure 21 indicates the effect of the Mk 1 
drag ring (pickle barrel) in reducing these localized 
high pressures. Although the drag ring, a light 
wooden structure used in standard service drops, 
was intended originally for stabilizing the airflight 
of the torpedo, experience at Newport has shown 
beneficial results in the reduction of damage to 
torpedoes. 


stage and continuing through the steady running 
phase. This instrument was designed to give roll, 
the angle of rotation of the torpedo about its longi- 
tudinal axis; pitch, the angle between the trajectory 
and the torpedo axis in a vertical plane; yaw, the 
similar horizontal angle; attitude, the angle between 
the longitudinal axis of the torpedo and any hori- 
zontal plane; and deviation, the angle formed by the 
intersection of a vertical plane through the longi- 
tudinal axis of the torpedo and the vertical plane 


36 


FACILITIES AND INSTRUMENTATION FOR STUDY OF TORPEDO ENTRY 


through a set course. The instrument consists of a 
Mk 12-1 gyro so modified that rotation between the 
outer gimbal ring and the torpedo and between 
inner and outer gimbal rings may be recorded. The 
instrument is contained in a cubical case, and the 
gyro is so oriented that the spin axis and the outer 
gimbal axis will always be released at 90 degrees to 
each other. The gyro is held by a centering pin 


inner and outer gimbals it is 0.75 degree. The step- 
wise record of contact closure obtained from the 
camera film is reduced to data of the type given on 
the roll record of Figure 22. Roll records of this 
type are much more satisfactory than those ob- 
tained with the Foxboro depth and roll recorder, 
because the gyro is not subject to inertia forces. 
The pendulum of the Foxboro instrument responds 


| 220 
J 200 

V) 

$ 180 

o ,40 

h 120 

o 100 
o 

80 

60 

40 

20 

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% 0 
o 20 
40 
60 
80 
100 
uj 120 

i"° 

O 160 

o 180 
200 
220 
240 






























































































































































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DLL ' 

VEUC 

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r 22 

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EG/! 

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/ 

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l 

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j 









V 





























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LAUNCHING NO. 1626 DUMMY MBB 

VELOCITY 334 FPS GROSS WEIGHT 1603 LB 

CG 0.53 IN. BELOW AXIS MINIMUM FINS 

INITIAL ANGLES 

ATTITUOE 18.4* YAW +3.6* 

PITCH 2.5* F ROLL -153* 


















































































































































3 4 

TIME IN SECONDS 


Figure 22. Roll-time record. No propellers. 


which can be disengaged by the starting accelera- 
tion of a tube launching or by a solenoid. 

The gyroscopic orientation recorder measures 
angle changes which are recorded on the neon tube 
camera in the torpedo by means of very light brushes 
sliding over commutators. The angular resolution 
of a camera commutator is determined by the 
spacing of the contacts and is 1 degree between the 
outer gimbal and the torpedo, whereas between the 


to centrifugal force which exists during any hooking 
or turning of the torpedo and gives a spurious 
indication of roll. Furthermore, the gyro instru- 
ment is not limited to the 30-degree travel of the 
Foxboro pendulum. 

Many other accessory instruments and measuring 
techniques were utilized in the project, but for such 
detail reference should be made to the general report 
on this project. 


PART II 


ROCKET PROPELLANTS AND INTERIOR BALLISTICS 

By B. H. Sage a 


D uring World War II, artillery rockets were 
again employed to advantage in a number of 
special tactical situations. This renewed interest 
in rockets may be ascribed in part to the greater 
mobility of arms and the consequent premium 
placed upon a low ratio of weight of launching 
equipment to weight of ammunition fired per unit 
time. The development of rockets for the U. S. 
Army and Navy was initiated in 1940 by the Na- 
tional Defense Research Committee. The dis- 
cussion in Part II summarizes the status of the 
interior ballistics of artillery rockets, their ignition, 
and the utilization of dry-processed double-base 
powders as propellants. There are also a few brief 
statements on the overall situation regarding the 
propulsion systems of such rockets and the probable 
future course of progress in this field. 

The material presented in Part II arises almost 
entirely from the development and experimental 
production activities of Section V of Contract 
OEMsr-418 between the Office of Scientific Research 
and Development and the California Institute of 
Technology. This work was carried out between 
October 1941 and October 1945, for the most part 
by the professional members of Section V. Primary 
emphasis was upon the designing, construction, 
testing, and semiproduction fabrication of relatively 
simple propulsion systems of artillery rockets . Little , 
if any, effort was made to achieve rockets of the 
highest performance, since the necessary meticulous 
refinement in design would have materially in- 
creased the time required for their development and 
decreased the number of rounds which could have 
been prepared with the limited facilities available. 
The marked emphasis which was placed upon the 

a Supervisor of Section V (Propellants and Interior Ballis- 
tics), Contract OEMsr-418, California Institute of Tech- 
nology. 


development and experimental production of spe- 
cific weapons prevented the systematic collection of 
as large a background of experimental facts con- 
cerning the underlying principles of interior bal- 
listics, ignition, and deflagration of double-base pro- 
pellants as would normally be expected in the 
course of a program of comparable scope carried out 
under less urgent conditions. Nevertheless, suffi- 
cient information has gradually been accumulated 
to permit a number of significant generalizations to 
be made, which are presented in some detail in two 
book-length publications. 1,2 

No further explicit reference will be made to these 
books, which in themselves represent a summary of 
the subjects under discussion and which serve to a 
large extent as the basis for the present limited 
treatment. On the other hand, specific references 
will be made whenever possible to the technical 
reports of Section V which contain the pertinent 
experimental data. Also all reports issued on the 
work of the section have been listed . b 

The technical progress which was realized by 
Section Y represents the efforts of at least 20 pro- 
fessional men. However, the work of W. N. Lacey 
and D. S. Clark of the staff of the California In- 
stitute of Technology was particularly helpful in 
connection with the supervision of certain of the 
activities. Acknowledgment should also be made of 
the significant contributions of R. N. Wimpress, 
W. H. Corcoran, and Q. Elliott to the field of 
interior ballistics and propellants, and of the assist- 
ance rendered by B. H. Levedahl and D. F. Botkin 
in the studies of physical and thermal properties, 
respectively. 


b In the general bibliography appended to this volume the 
principal Section V reports are listed under OEMsr-418 
designations IAC, IBC, ICC, IDC, IGC, JAC, JBC, JCC, 
JDC, and JGC. 



37 









• • 























Chapter 5 

INTERIOR BALLISTICS 

By B. H. Sage 


51 PRINCIPLES OF ROCKET 

PROPULSION* 

T he principles of rocketry, which are rela- 
tively simple, have been known for an extended 
period of time. In a general way, the relationship 
between the exterior ballistic behavior of the round 
and the performance of the rocket motor can be 
indicated in the following manner for a projectile 
containing a weight W of inert components and a 
weight w of propellant traveling at a velocity V 
with a thrust F applied. Under these circumstances, 
using t for time and g for acceleration of gravity, 
the acceleration is given by 


The weight of propellant changes as burning pro- 
gresses and the products of combustion are expelled 
through the nozzle. If air drag and other minor 
effects are neglected, the velocity at the end of 
burning, F 0 , hereafter called the burnt velocity, 
may be evaluated by the following expression, 
where wo is the weight of the propellant and IF is 
the weight of the inert parts of the round: 


As a matter of interest, values of effective gas 
velocity for a number of the common rocket pro- 
pellant combinations used in this country are pre- 
sented in Table 1. These values are applicable only 
to the specific combinations of propellant and metal 
parts indicated. 

One of the more effective measures of the overall 
efficiency of a rocket motor is the specific impulse, 
that is, the impulse available per unit total weight 
of propellant or of motor. In general, a well-de- 
signed rocket motor should yield an overall impulse 
of approximately 100 lb-sec per lb of motor; but 
not many of the rocket motors developed during 
World War II gave such high performance. Values 
of impulse per total unit weight of rocket motor for 
a number of the common rocket motors are included 
in Table 1. However, experimental work now in 
progress b under the cognizance of the Navy in- 
dicates that rocket motors can be developed with 
an overall specific impulse of approximately 120 lb- 
sec per lb. Such improvements result from careful 
revisions in design so as to decrease to a minimum 
the weight of the metal parts as compared to that 
of the propellant. 


< 2 > 

For convenience, it is desirable to relate the thrust 
and the weight rate of burning of the propellant, 
dw/dty by means of a term called the effective gas 
velocity, which is essentially constant for a given 
propellant burning in a particular rocket motor: 

f = y 4r. < 3 > 

A combination of equations (2) and (3) results in 
the simplified expression: 

Vo = V E In (l + (4) 


a See Parts III and V for different treatments of these 
principles. 


Table 1. Performance of JPN* propellant in several 
rocket motors. 


Rocket 

motor 

Grain f 

Temp. 

(°F) 

Effective 

gas 

velocity 

(fps) 

Specific impulse 
(lb-sec per lb) 
Propellant Rocket 
= Vs/g motor 

2.25-in. Mk 10 

Mk 1 

10 

6,670 

207 

32.6 



70 

6,600 

205 

32.3 



130 

6,350 

197 

31.0 

3.25-in. Mk 6 

Mk 13 

0 

6,180 

193 

53.5 



70 

6,900 

205 

56.9 



140 

6,420 

202 

56.0 

5.0-in. Mk 1 

Mk 18 

-20 

6,600 

206 

57.2 



0 

6,680 

207 

57.4 



70 

6,930 

215 

59.7 



140 

7,000 

218 

60.5 



160 

6,830 

212 

58.8 


* See Table 2 for composition, etc. 
t See Table 4 for ballistic characteristics. 


b In summer of 1946. 


39 



40 


INTERIOR BALLISTICS 


5 2 PRACTICAL LIMITATIONS 

The performance of rocket motors is limited by 
a number of practical considerations. The change 
in enthalpy upon reaction of the propellant does not 
in most instances exceed approximately 2,300 Btu 
per lb. Maximum specific impulse for a JPN pro- 
pellant is presented in Figure 1 as a function of the 
reaction pressure and the expansion ratio of the 
nozzle. The performance indicated is the maximum 


since the increase in pressure at the front end of the 
rocket motor over that obtaining at the nozzle 
causes a more rapid increase in burning rate than is 
compensated for by the increased flow arising from 
the higher pressure differential. In addition, rela- 
tively long grains fail as columns near the end of 
burning and either prevent the egress of gas from 
the motor, with a consequent failure of the metal 
parts, or yield large losses of unburned propellant, 
with a corresponding decrease in specific impulse. 



Figure 1 . Possible specific impulse as function of reaction pressure and change in enthalpy. 


that may be obtained; hence the effective impulse 
in actual operations is less, depending upon the loss 
of unburned propellant, frictional effects, and other 
causes. 

Since the specific weight of most propellants is of 
the order of 100 lb per cu ft, the quantity of ma- 
terial to be placed in any given cross section of 
round is limited. Therefore, it is only possible to 
increase the quantity of propellant in a round of 
given cross section by increasing the length . 

It is not possible, however, to increase the 
length indefinitely in the case of rockets with the 
nozzles located at a single section along the round, 
inasmuch as, when the burning occurs normal to the 
axis of the grain, frictional effects become of in- 
creasing importance as the round is lengthened. 
These limitations finally become controlling, and 
the reaction of the propellant becomes unstable, 


These practical limitations can be overcome to a 
certain extent by the use of nozzles located at 
several sections along the axis of the rocket motor, 
but the resulting added complexity does not appear 
to justify this procedure except in a few special 
cases. Moreover, rounds which are excessively long 
in comparison to their diameter usually constitute 
a difficult handling problem. In general, it does not 
appear advantageous to utilize rocket motors whose 
length is much greater than 12 times the caliber of 
the round. 

It may be of interest to note that liquid propel- 
lants do not impose the restrictions upon the 
geometry of the rocket motor that are encountered 
in the case of rockets with solid fuel. When liquid 
propellants are used, the reaction chamber may be 
made relatively small, and the fuel may be stored 
in containers of any shape suited to the exterior 



BURNING CHARACTERISTICS OF PROPELLANTS 


41 


ballistic requirements of the round. The stowage of 
such artillery rockets, however, constitutes a prob- 
lem that has not yet been solved. Nevertheless, it is 
believed that the use of liquid fuel in the larger 
artillery rockets is well worth consideration. The 
German government realized some success with 
liquid-fueled rockets, which in many instances ex- 
hibited superior ballistic characteristics to the solid- 
fueled rockets of comparable caliber. However, 
stowage difficulties were often encountered because 
of the corrosive action of the fuel. 


5 3 BURNING CHARACTERISTICS OF 
PROPELLANTS 

The colloidal double-base dry-extruded fuels used 
in artillery rockets burn upon the exposed surfaces 
at a weight rate which is roughly proportional to 
the area exposed. It is therefore of importance in 
the design of charges for artillery rockets to ensure 
that the change in burning area as the reaction 
proceeds is in accordance with the ballistic require- 
ments imposed. For example, a marked change 
in the weight rate of reaction can be realized by 
relatively small changes in the cross section of 
the round. The reaction pressure also exerts a sig- 
nificant influence on the burning rate, as does the 
temperature of the propellant. 

5,31 Influence of Position in Grain 
upon Burning Rate 

It has been shown by numerous experiments that 
the burning rate of a solid propellant increases as 
the center of the web is approached. This is prob- 
ably due in part to the gradual increase in the 
temperature of the unburned propellant because of 
thermal transfer and in part to the somewhat higher 
rate of transfer of radiant energy from the motor 
walls, the temperature of which rises during the 
latter part of burning. However, experimental 
measurements indicate that, even when the radia- 
tion and temperature effects described above have 
been eliminated, the burning rate of propellant 
under particular conditions of pressure and tem- 
perature is higher near the center of the web than 
near the original surfaces of the grain. 1 In Figure 2 
is shown the influence of position upon burning rate 
for JP propellant, whose composition is given in 


Table 2, together with that of other typical propel- 
lants. It is apparent that the influence of position 
is significant in that the final burning rate for a 
propellant temperature of 0 F is higher than the 
initial burning rate for a propellant temperature 
of 70 F. 



Figure 2. Instantaneous burning rates for JP 
propellant, showing influence of position. 


532 Influence of Gas Velocity 

There is, in addition, a significant effect which is 
directly related to the flow of the products of reac- 
tion past the reacting surface. In the case of rela- 
tively high weight rates of flow, the burning rate 
may be 30 or 40 per cent higher in the region of high 
gas velocity than where the reaction surface is sur- 
rounded by an essentially stagnant gas phase. This 
influence of erosion is shown in Figure 3, which pre- 
sents comparative photographs of partially burned 
grains at the front and nozzle ends. 


42 


INTERIOR BALLISTICS 


5 3 3 Influence of Pressure 

For most double-base propellants, the burning 
rate is significantly influenced by the pressure of 
the reaction. This may in part result from the 
higher rate of energy transfer to the propellant 
from the products of reaction by radiation at the 
higher pressures. The influence of pressure may be 
approximated by an exponential equation of the 
general form 

B = ^(1,000) ’ (5 ^ 


tures indicated. The table also records values of the 
exponent n. 

534 Influence of Temperature 

Temperature also influences the burning rate of 
a propellant. 1 The burning rate of the propellants 
which are markedly influenced by the reaction 
pressure are also particularly susceptible to the 
changes in the temperature of the charge. The 
importance of decreasing these effects is difficult to 
exaggerate since one of the primary attributes of 



Figure 3. Partially burned grains showing appreciably less erosion at the front end (A) than at the nozzle 
end (B). 


where (3 and n are single-valued constants of the 
initial temperature of the propellant. The term 
p'/ 1,000 is used in place of the pressure in order 
to simplify numerical calculations and permit ready 
comparison of burning rates. It should be realized 
that at a reaction pressure of 1 ,000 psi abs the con- 
stant j8 represents the burning rate. Values of this 
burning rate at 1,000 psi for several common pro- 
pellants are presented in Table 3. It should be 
emphasized that these values represent instantane- 
ous burning rates at the initial propellant tempera- 


a good rocket propellant is the small influence 
which reaction pressure and propellant temperature 
have upon the burning rate. Although most 
existing colloidal propellants exhibit relatively 
large variations in burning rate with temperature, 
it is evident from Table 3 that for certain of these 
the effect is much smaller than for the others. It is 
believed, therefore, that significant advancement 
in this direction can be made by careful investiga- 
tion of the propellants which show the smaller 
influences of pressure and temperature upon burn- 


BURNING CHARACTERISTICS OF PROPELLANTS 


43 


Table 2. Composition and some thermal properties of typical rocket propellants. 


Identification 

Type 

Composition 

Constituent 

Weight 
(per cent) 

Heat of 
explosion 
(cal per gm)* 

Adiabatic 

flame 

temperature 

(°F)f 

JP 

Ballistite of same com- 

Nitrocellulose (13.25 per cent N) 

52.2 

1230 

5300 


position as trench mor- 

Nitroglycerin 

43.0 




tar sheet powder 

Diethylphthalate 

3.0 





Diphenylamine 

0.6 





Potassium nitrate 

1.25 





Nigrosine dye (added) 

0.1 



JPN 

Ballistite, modification 

Nitrocellulose (13.25 per cent N) 

51.5 

1230 

5300 


of JP formula to imr 

Nitroglycerin 

43.0 




prove stability 

Diethylphthalate 

3.25 





Ethyl centralite 

1.0 





Potassium sulfate 

1.25 





Carbon black (added) « 

0.2 





Candelilla wax (added) 

0.08 



JPH 

Powder with burning 

Nitrocellulose (12.6 per cent N) 

55.5 

1260 

5450 

(FDAP 60) t 

properties similar to 

Nitroglycerin 

42.0 




JPN, but of higher 

Ethyl centralite 

1.0 




physical strength 

Potassium sulfate 

1.5 





Carbon black (added) 

0.2 





Candelilla wax (added) 

0.02 



Russian cordite 

Double-base powder, 

Nitrocellulose (12.2 per cent N) 

56.5 

880 

3750 

(FDAP 44) t 

cooler and slower burn- 

Nitroglycerin 

28.0 




ing than JPN 

Dinitrotoluene 

11.0 





Ethyl centralite 

4.5 





Candelilla wax (added) 

0.08 



H-4§ 

Double-base powder 

Nitrocellulose (13.15 per cent N) 

58.0 

950 

4000 


with burning rate inter- 

Nitroglycerin 

30.0 




mediate between that 

Dinitrotoluene 

2.5 




of JPN and that of Rus- 

Ethyl centralite 

8.0 




sian cordite 

Potassium sulfate 

1.5 





Carbon black 

0.02 



Ball powder 

Compression-molded 

Nitrocellulose (12.5 per cent N) 

45.3 

1050 

4500 


powder made by West- 

Nitroglycerin 

45.0 




ern Cartridge Co. 

Trinitrotoluene 

9.0 





Ethyl centralite 

0.7 



218B 

Compression-molded 

Ammonium picrate 

46.5 




composite propellant 

Sodium nitrate 

46.5 





Butyl ureaformaldehyde resin 

5.1 





Plasticizer 

1.5 





Calcium stearate 

0.4 




* Heat of explosion at constant volume with water in reaction products as liquid, 
t Temperature with reaction at constant pressure. 

t These numbers identify experimental lots of propellant manufactured by the Sunflower Ordnance Works, Lawrence, Kansas, that arc representative 
of the designated types. 

§ Designated T-2 by the Ordnance Department. 



44 


INTERIOR BALLISTICS 


Table 3. Average burning rate data for typical propellants. 


Powder 

n* 

0(ips)* 

(S In 

V ST ) v 

/ d In p\ ft 

v aT A.v 



(0 F) 

(70 F) 

(140 F) 

(1/°F) 

(1/°F) 

JP 

0.71 

0.551 

0.671 

0.815 

0.0028 

0.0096 

JPN 

0.69 

0.564 

0.651 

0.752 

0.0021 

0.0068 

H-4f 

0.65 

0.330 

0.380 

0.437 

0.0020 

0.0057 

Russian (FDAP 44) 

0.70 

0.250 

0.290 

0.337 

0.0021 

0.0070 

German 

0.71 

0.188 

0.218 

0.254 

0.0022 

0.0076 

Japanese 

0.42 

0.278 

0.311 

0.349 

0.0016 

0.0028 

Western Cartridge 

0.64 

0.340 

0.393 

0.454 

0.0021 

0.0058 

218B composite 

0.52 

0.700 

0.750 

0.802 

0.0010 

0.0021 

JPH 

0.69 

0.581 

0.676 

0.785 

0.0022 

0.0071 


* Constants to use in relation B = (3 (p'/l,000) re . f Designated T-2 by the Ordnance Department, ft Kjf, nozzle coefficient, is the ratio of burning 
area to nozzle throat area. 


ing rate; that is, the propellants of the H-4 type c 
rather than the JP or JPN types. 

5 4 OPTIMUM PROPELLANT * 

CHARACTERISTICS 

Sufficient experience has now been accumulated 
to indicate the characteristics which are particularly 
desirable in a propellant for use in artillery rockets. 
It should be realized that these so-called optimum 
characteristics are from necessity somewhat general 
and that the importance of each of the several 
factors differs widely with various applications. 

541 Burning Rate 

It is desirable that the influence of pressure, 
temperature, radiation, and gas velocity upon burn- 
ing rate be as small as is feasible. 1-3 Such a propel- 
lant can probably be approached by making suitable 
adjustments in composition and providing for ade- 
quate opaqueness. Modifications of composition 
are particularly efficacious with propellants of 
somewhat lower potential than the JP group — the 
H-4 stock, for example. 

Physical Properties 

The propellant should have adequate compressive 
strength and be resistant to impact, especially at 

c Developed at Allegany Ballistics Laboratory for the 115- 
mm aircraft rocket. This propellant composition is designated 
T-2 by the Army Ordnance Department. See Chapter 13 of 
this volume. 


low temperatures. 4-11 It appears that ultimate 
compressive strength is an index of the performance 
of- colloidal propellants under conditions where great 
axial stress is applied to the grain during deflagra- 
tion; and high impact values are important in the 
handling of rocket motors, especially at low tem- 
perature, since malperformance may result if the 
impact energies are not at least comparable to those 
realized with JPN propellant (approximately 12 ft- 
lb per sq in. at a temperature of 0 F). It is also 
necessary that the compressive strength not deteri- 
orate unduly at the higher temperatures. For ex- 
ample, unsatisfactory field performance is obtained 
at 140 F with the Mk 13 grain in the 3.25-in. 
rocket motor Mk 6 when JP propellant, which has 
an ultimate compressive strength (at this tempera- 
ture) of approximately 270 psi, is used. 8 However, 
satisfactory performance may be obtained with the 
same round at temperatures up to 150 F by using 
a propellant of identical ballistic characteristics but 
an ultimate compressive strength of 1,300 psi at 
140 F. 7 

543 Stability 

Although reasonable chemical stability is an im- 
portant characteristic, most propellants involve 
compounds that tend to decompose with time. Ni- 
trocellulose is especially troublesome in this regard, 
since its stability is significantly affected by manu- 
facturing techniques and its rate of decomposition 
cannot usually be predicted with certainty. In 
order to improve colloidal double-base propellants 
from this standpoint, it will be necessary to investi- 
gate the characteristics of nitrocellulose, with par- 
ticular attention to the influence of manufacturing 


EFFECTS OF ACCELERATION 


45 


techniques and the nature of the cellulose employed. 
In addition, the investigation should include suit- 
able stabilizers; for the principal improvement of 
JPN over JP powder is in stability, which is 
apparently attributable to the substitution of ethyl 
centralite for diphenylamine. d 

It is also important that the propellant be of such 
a physical nature as to be geometrically stable with 
respect to time. Any significant change in the geom- 
etry of the grain during stowage will result in a 
corresponding change in ballistic characteristics. 
These changes may be extensive enough to cause 
failure of the round. 


544 Toxicity 

It is important from a processing and loading 
standpoint that the propellant be as nontoxic as is 
compatible with satisfactory performance. Low 
toxicity is not a controlling requirement but is cer- 
tainly a desirable attribute if the propellant is to be 
manufactured in large quantities with a minimum 
of special equipment and the fewest possible physi- 
ological difficulties for the operators. 


Specific Impulse 

The specific impulse of the propellant should be 
the highest that is feasible. In this respect, pro- 
pellants now vary from approximately 100 to 220 
lb-sec per lb; and it does not appear that many will 
be found in the near future for which the specific 
impulse will exceed the latter value. 


5 5 INFLUENCE OF BURNING TIME 
ON TOTAL IMPULSE 


From the standpoint of exterior ballistics, it is 
usually desirable that the burning time be as short 
as feasible, since for most types of rockets dispersion 
increases with burning time. However, this factor 
is not of great importance in connection with for- 
ward-firing fin-stabilized rockets launched from the 


d For further information on stabilizers and their effects on 
propellant characteristics, see sections of the Division 8 
Summary Technical Report covering the work of Pauling at 
CIT under Contract OEMsr-881, and reports submitted under 
that contract. 


exterior of aircraft, because the airstream induces 
an inherent stability in the rocket at the time of 
launching. The design of a rocket is therefore a com- 
promise between the requirements which must be 
met in order to obtain low dispersion and high 
impact velocities at short ranges, and the limita- 
tions which are imposed by the design of the charge . 

The weight of propellant per unit cross section of 
rocket is roughly a function of the burning time. 
The influence of burning time on the specific im- 
pulse per unit cross section of the round is shown in 
Figure 4 for JPN powder. e This relationship is not 



BURNING TIME IN SECONDS 


Figure 4. Impulse per unit cross-sectional area 
as a function of burning time for three types of 
charges. 


strictly single-valued but covers a wide range, 
depending upon the particular grain section em- 
ployed. Ultimately, the optimum behavior would 
be obtained with grain burning only on one end; 
but the burning time with existing propellants 
would be unduly long. 


56 EFFECTS OF ACCELERATION 

Acceleration imposes relatively large setback 
forces upon propellant grains. In the case of the 

e The curves shown are based upon a burning rate of 0.65 
ips, an internal area ratio (of burning area to ports area, i.e., 
the cross section available for gas flow) of 100 for tubular 
grains, and a motor of 5-in. inside diameter. 



46 


INTERIOR BALLISTICS 


Mk 13 grain, for example, 11,12 a total force of ap- 
proximately 340 lb is applied to the grid f by the 
grain during the early part of the acceleration of a 
round fired at 70 F, or about 525 lb for a round 
fired at 120 F. These forces cause elastic, and under 
some conditions plastic, deformation of the grain 
near the nozzle, with a corresponding decrease in 
the cross-sectional area through which the products 
of combustion flow from the forward end of the 
rocket motor to the nozzle. Since such port areas 
are relatively critical near the upper operating tem- 
perature limit of the round, relatively small changes 
in port area resulting from the elastic and plastic 
deformation of the grain may influence significantly 
the temperature at which unstable burning occurs. 

Near the end of burning, the slenderness ratio of 
a grain becomes much larger; and, although the 
total force attributable to acceleration is smaller, 
the force per unit area resulting from the accelera- 
tion may nevertheless be enough to cause breakup of 
the grain in flight when practically no disintegra- 
tion would occur under static conditions. If suffi- 
ciently extensive, the breakup of the grain will 
result in the failure of the round because of the 
marked increase in burning area. In any event, it 
will cause a distinct increase in pressure and the 
loss of unburned propellant. 



Figure 5. Pressure-time curves for Mk 13 grains 
showing breakup near end of burning at 140 F. 


A significant part of the work of designing charges 
for rocket motors has involved studies of the in- 
fluence of composition on physical character- 
istics 9,11,13 of propellants in order to decrease not 
only breakup near the end of burning but also 
deformation at the beginning. The quantitative 
nature of the breakup of grains is shown in Figure 5 
and is described in some detail elsewhere. 12,14 

f A rocket component, usually of steel, which supports the 
rear end of the grain, and is supported by the nozzle. 


57 TEMPERATURE LIMITS 

The operating temperatures of rocket ordnance 
are greatly limited by the effects of temperature 
upon the physical and chemical characteristics of 
the propellant. At low temperatures the propellant 
becomes more brittle; 7,9 consequently, the grain 
may fail as the result of stresses imposed by accel- 
eration or accidental impacts encountered in han- 
dling. Furthermore, at low temperatures the burn- 
ing rate of the propellant decreases sufficiently for 
unstable burning to occur, during which the reaction 
substantially ceases and the propellant is reignited 
after an interruption of as long as a second or two. 
The reignition may be caused by contact with the 
hot metal parts of the rocket. In general, small 
grains begin to show unstable burning when the 
reaction pressure falls below 400 psi. However, in 
the case of large grains, where there are usually 
somewhat thicker gas films and where the energy 
loss per unit weight of propellant is somewhat 
smaller, stable reactions can be maintained at 
much lower pressures. 

With respect to reaction pressure, the lower limit 
of stability depends upon the geometry of the 
particular charge under consideration. 15-18 At high 
temperatures the increase in burning rate introduces 
marked increases in pressure within the reaction 
chamber as a whole, and in many instances the 
upper limit of propellant temperature at which the 
round may be successfully fired is determined by the 
maintenance of stable burning near the end of the 
round opposite that at which the nozzles are 
located. Instability is not often encountered with 
rounds for which the ratio of burning area to port 
area is less than 100; however, the weight of propel- 
lant which may be stored in each unit cross-sectional 
area is limited . As a matter of interest , a number of 
the more pertinent interior ballistic characteristics 
of the several principal rocket charges developed by 
OEMsr-418 during World War II are recorded in 
Table 4. 

For rounds in which the upper temperature limit 
is not controlled by the pressure developed during 
the reaction or by the occurrence of unstable burn- 
ing of the propellant, it is limited by the physical 
properties of the propellant. The ultimate com- 
pressive strength decreases markedly with increase 
in temperature, 7,8,10 and in each case a temperature 
is reached at which the charge will not withstand the 
acceleration and frictional forces without under- 


CHARGE DESIGN 


47 


going significant plastic deformation. Under these 
circumstances the port area is decreased, and even- 
tually a condition of unstable burning is reached. 


Table 4. Ballistic characteristics of several rocket 
motor charges. 


Grain 

Mk 1* 

Mk 13f 

Mk 18J 

Burning area, A c (sq in.) 

Initial 

98.9 

281.4 

598 

Final 

66.4 

260.0 

613 

Free port area, A v (sq in.) 

Initial 

0.96 

2.54 

6.3 

Final 

3.14 

6.90 

16.8 

Ratio of burning area to 
port area, A c /A v 

Initial 

103 

110.7 

105 

Final 

21 

37.7 

36.5 

Average nozzle pressure (psi) 

-20 F 


340 

610 

0 


450 

734 

20 

874 



70 

1,580 

850 

1,071 

130 

2,587 


140 


1,330 

1,902 


* In 2.25-in. Rocket Motor Mk 9. 
t In 3.25-in. Rocket Motor Mk 7. 
t In 5.0-in. Rocket Motor Mk 1. 


The influence of temperature on the ultimate com- 
pressive strength of two propellants is presented 
graphically in Figure 6, which also shows the cor- 
responding stresses imposed during firing under both 
static and flight conditions for the 3.25-in. rocket 
motor with a Mk 13 grain. 

58 CHARGE DESIGN 

The design of propellant charges for rocket motors 
is controlled by the exterior ballistic requirements of 
the rocket and the deflagrating characteristics of the 
propellant. Since the object is usually to obtain a 
relatively uniform acceleration, which involves a 
constant weight rate of discharge from the nozzle 
during deflagration, 1, 19-22 it is customary as a first 
approximation to design charges to burn neutrally. 
However, because of the thermal energy transferred 
to the metal parts of the motor and the consequent 
decrease in tensile strength of these parts, it may be 
necessary to arrange for the reaction pressure to 
decrease as the reaction proceeds. 

This regressive type of charge design is particu- 
larly desirable in external-burning grains, such as 
the cruciform, where large changes in the physical 


characteristics of the metal parts may take place 
during the burning interval. In addition, there is a 
significant erosion of most nozzles, especially when 
the reaction pressure is high and the nozzle diameter 
small . This tends to increase the weight rate of flow 
for a fixed reaction pressure. Therefore, even if the 



Figure 6. Influence of temperature upon the ul- 
timate compressive strength of JPH and JPN pro- 
pellants. 


charge is neutral in so far as its geometry is con- 
cerned, there will be a regression in the pressure as 
the reaction proceeds. For this reason it is often 
possible to design a neutral-burning grain and 
obtain the advantages of regressive burning by an 
increase in nozzle area from erosion. On the other 
hand, it is impossible to lengthen a particular rocket 
grain indefinitely without reaching an unstable 
situation wherein the increase in burning rate re- 
sulting from the rise in pressure is greater than the 
increase in rate of flow resulting from the same rise 
in pressure. 

A number of typical grain sections employed in 
rockets developed during World War II or under 
investigation at that time are illustrated in Figure 7 . 
All these charges burn externally, or both internally 
and externally, and hence require sufficiently heavy 
metal parts to withstand the reaction pressure at 
the end of burning, when the average temperature 


48 


INTERIOR BALLISTICS 


of the metal parts is markedly higher than at the 
beginning. It appears that a significant decrease in 
the weight of the metal parts may be realized by 
utilizing an internal-burning grain of the type 
shown in Figure 8, which is inhibited on the periph- 
ery to prevent burning except in the axial per- 
foration. 14,23-26 More recent experience with this 








Figure 7. Cross sections of typical grains. 

type of grain than was covered by the work of 
Section V of OEMsr-418 indicates that the surface 
temperature of the metal parts may be maintained 
below 140 F at all points except where the products 
of reaction come in contact with the interior of the 
wall. Such a grain is particularly adaptable for use 
in spin-stabilized rockets since it is well supported 
against centrifugal forces. 

Internal-burning grains are considered one of the 
most promising means of increasing the performance 
of rocket motors, for it has been shown that their 
use will permit the overall specific impulse of the 


motor to be raised to a value well above 100 lb-sec 
per lb. It is possible to combine the internal-burn- 
ing grain with the external-burning grain to obtain 
a relatively high density of loading and yet keep the 
burning time within the limits imposed by the 
exterior ballistic requirements for many types of 
rounds.® 

The design of a solid fuel propulsion system for a 
particular application is based upon the funda- 
mental principles of interior ballistics, 21,22,27 as well 
as upon an appraisal of the desired characteristics 
of the round. For example, if the round is to be one 
of maximum burnt velocity and straight trajectory, 
such as an antiaircraft rocket, it is necessary to 



Figure 8. Internal-burning grain with cog- 
shaped axial perforation. 


obtain the maximum impulse per unit of cross 
section. An end-burning grain of sufficiently rapid 
burning rate to give the desired acceleration and 
maintain a high terminal velocity would probably 
meet the requirements. However, since existing 
propellants do not even approach the necessary 
burning rate, an interior-burning grain must be 
employed . 

The length of the grain which is to be used, and 
hence the average weight per unit cross section, is 
limited by the ratio of the burning area to the port 
area. This ratio is presented in Figure 9 as a func- 
tion of the percentage of the cross-sectional area 
occupied by propellant for several ratios of the 
length to the diameter of an internal-burning grain 
with cylindrical axial perforation. This represents 
the minimum value of the ratio of burning area to 

« This discussion does not include multiple-grain charges of 
the conventional type, such as were widely used with solvent- 
processed double-base propellants and in a number of foreign 
rockets. Such charges appear to have been effective in a num- 
ber of applications; but they preclude insulating the motor 
wall from the transfer of thermal energy with the grain itself 
as can be done in the case of the internal-burning charge. 


RECOMMENDATIONS 


49 


port area that may be obtained with any shape of 
interior perforation and, therefore, is the optimum 
fraction of the cross-sectional area that may be 
occupied by propellant. However, the cylindrical 
cross section cannot usually be employed, since it 



20 40 60 80 


PER CENT OF CROSS SECTION OCCUPIED BY PROPELLANT 

Figure 9. Influence of size of cylindrical perfora- 
tion upon characteristics of an internal-burning 
grain. 

results in an unduly progressive charge, except in 
instances where the web thickness is so small that an 
axial perforation with irregular periphery is not 
required. 



0 0.20 0;40 0.60 0.80 1.00 


RATIO OF PORT AREA TO CROSS-SECTIONAL AREA 
OF ROUND 

Figure 10. Influence of relative port area in an 
internal-burning grain upon loading density. 

The data presented in Figure 9 permit the evalu- 
ation of the ratio of port area to cross-sectional area 
that should be employed in order to obtain the 


maximum weight of propellant in a particular rocket 
motor. In Figure 10 the weight of propellant that 
can be loaded into each unit cross-sectional area of 
a motor is shown as a function of the ratio of port 
area to cross-sectional area. In this instance it is 
assumed that the maximum acceptable ratio of 
nozzle port area to burning area is 100. 

59 RECOMMENDATIONS 

In the opinion of the writer, the use of internal- 
burning grains is the most promising approach to 
future developments in charge design and interior 
ballistics of rockets using solid fuels. In the case of 
short grains in which burning time is not important, 
a relatively small port area may be employed with a 
corresponding increase in the weight of propellant 
per unit of length and cross section. In situations 
where burning time is of importance, a propellant 
grain of the cross section illustrated in Figure 11 



Figure 11. A shell-and-rod charge extruded as a 
single grain. 

should prove useful. This grain is shown as ex- 
truded in a single piece, with the burning taking 
place on each of the exposed surfaces except the 
periphery. By appropriate modification of this 
design, it should be possible to obtain almost any 
desired burning time for a grain of given cross 
section. 11 


h The development of extrusion techniques and the details 
of the design of such grains have been carried out at the Naval 
Ordnance Test Station, Inyokern, subsequent to the termina- 
tion of active work under OEMsr-418. 



50 


INTERIOR BALLISTICS 


More specifically, the use of internal-burning 
grains in a number of applications appears desirable 
for the following reasons: 

1 . This type of grain avoids heat transfer to most 
of the wall of the rocket motor, thereby permitting 
the use of tubes of thinner steel, or possibly alu- 
minum alloy, with a corresponding increase in the 
specific impulse of the motor as a whole. 

2. The extrusion of concentric-web charges as a 
single unit will permit the production of internal- 
burning grains which are relatively simple to load 
and will withstand the high radial stresses asso- 
ciated with spin-stabilized rockets and still yield 
short burning times. 

3. The internal-burning grain, or a variation 
thereof, permits nearly the optimum quantity of 
propellant per unit cross-sectional area that can be 
obtained with any geometric design yet pro- 
posed. 

4. Present information indicates that internal- 
burning grains of JPN propellant burn stably over a 
relatively wide range of conditions and may be 
ignited without difficulty. 

5. It is not necessary to provide a conventional 
grid for these grains. 

6. The inhibiting of the exterior of the grains does 
not appear to constitute a production problem and 
may be accomplished by the application of cellulose 
acetate or ethyl cellulose as a spirally wrapped strip, 
a flat wrapped sheet, or a hot molded envelope. 

It is believed that by the use of internal-burning 
grains rocket motors can be constructed to give 
overall specific impulses significantly in excess of 
100 lb-sec per lb. Work should be directed toward 
the investigation of internal-burning charges which 
are closed at the end of the grain away from the 
nozzle, thus avoiding heat transfer to metal parts 
except in the immediate vicinity of the nozzle. In 
the case of long-range artillery rockets, this type of 
grain might be supplemented by an end-burning 
charge which would supply sufficient thrust to main- 
tain high velocity after the end of burning of the 
primary charge. Such grains could be inhibited in a 
single piece. It should be emphasized, however, that 
these latter recommendations have not yet been 
investigated and hence should be considered only as 
proposals for future study. 

At the present time the development of several 
types of internal-burning grains is in progress under 
the supervision of the Services. These should be 


useful in both spin- and fin-stabilized rockets. In 
applications where long burning time is permissible, 
a single axial perforation will probably suffice ex- 
cept in units of exceedingly large diameter. 

5 10 LIQUID FUELS 

As has been indicated, solid fuels have a number 
of limitations, notably the significant influence of 
temperature upon the ballistic and physical char- 
acteristics of the propellant. Moreover, the Ger- 
mans had notable success with the use of liquid 
fuels in at least one large guided missile and in a 
limited number of simpler artillery rockets. It is 
believed, therefore, that the use of liquid fuels in 
large artillery rockets should be given careful con- 
sideration. Solid fuel may be used as a pressurizing 
agent, and the fuel containers need only be designed 
to withstand the reaction pressure at ambient tem- 
perature. The reaction chamber may be relatively 
light, and film cooling may be employed. 

One of the primary requirements for a satisfactory 
liquid fuel for an artillery rocket is stability. At the 
present writing, binary liquid propellants seem to be 
more desirable for large artillery rockets than mono- 
liquid propellants. The probability of the detona- 
tion of a binary liquid propellant by small arms fire, 
or even high explosives, is small, whereas there will 
nearly always exist an energy threshold above which 
a mono-propellant will detonate. It appears that 
liquid-fueled rockets could be constructed in the 
larger sizes with a higher specific impulse for the 
rocket motor as a whole than the corresponding 
solid-fueled rockets. The cost of the metal parts 
may be somewhat higher; but, if the binary com- 
bination which is chosen shows adequate stability, 
the increase in performance would probably justify 
the added expense. 

The transition from solid- to liquid-fueled rockets 
should be considered at calibers between 8 and 14 
in., in so far as can now be determined. It does not 
seem practical to prepare single-grain solid fuel 
charges in diameters larger than perhaps 12 in. 
On the other hand, the use of liquid fuels in small 
rockets appears to be an unwarranted complication. 
The actual sizes and applications in which these 
two types of rockets will prove respectively 
superior remain to be established by develop- 
ment and Service experience.* At the present time 


LIQUID FUELS 


51 


it is believed that the use of the oxides of nitrogen 
or nitric acid as the oxidant and aniline or one of its 
derivatives as the fuel is the most promising com- 
bination for the immediate development of liquid- 
fueled rockets. Hydrogen peroxide-hydrazene hy- 
drate combinations do not appear well adapted to 
artillery rockets because of the difficulty of extended 
storage of hydrogen peroxide in sealed metal 
containers. 


In conclusion it is reiterated that the develop- 
ment of liquid-fueled artillery rockets utilizing 
binary spontaneously ignitable liquid propellants 
appears to be worth while in spite of the added 
hazards involved, because of the marked simplifica- 
tion in the ignition system. This opinion is based 
upon satisfactory experience with colloidal pro- 
pellants, which nearly always ignite if the case of 
the rocket motor is penetrated by gunfire. 


Chapter 6 

IGNITION 

By B. H. Sage 


6i GENERAL PRINCIPLES 

I gnition in rockets has for the most part been 
satisfactorily accomplished with black powder 
igniters initiated by electric squibs, although in a 
number of instances percussion units have been 
employed. Other types of igniter charges have been 
investigated at least to some extent; but, of these, 
organic materials such as double-base propellants 1 
have not proved particularly successful, and metal- 
oxidant mixtures, 2,3 although acceptable from the 
ballistic standpoint, offered no significant advantage 
over black powder in this respect and at the same 
time appeared to be somewhat more hazardous to 
handle. The mixture of this type which was found 
most satisfactory, magnesium powder and potas- 
sium perchlorate, is subject to detonation when 
fired in significant quantities; hence no extensive 
investigation was made of its detailed application. 
Moreover, since a large number of munition manu- 
facturers are familiar with the methods of process- 
ing black powder, it is believed that the continua- 
tion of its use as an igniter charge in rockets fueled 
with double-base propellants is desirable. The 
present discussion will therefore be confined to 
black powder igniters and their characteristics. 

In principle the ignition of a rocket motor utiliz- 
ing a double-base powder as the propellant consists 
in transferring energy to the propellant at a suffi- 
ciently high rate to bring the immediate surface 
to the autoignition temperature, which is approx- 
imately 340 F. The detailed mechanism associated 
with this process is not well understood, although it 
appears that the igniters in question function pri- 
marily by the radiant transfer of energy from the 
products of reaction of the black powder to the pro- 
pellant. Since the products of reaction of the black 
powder have a somewhat higher emissivity than 
those of the propellant, unusually high rates of 
burning of the propellant are obtained during the 
period that the products of reaction of the igniter 
are within the reaction chamber. This behavior is 
illustrated by Figure 1, which shows pressure as a 
function of time for the Mk 18 grain at several 




Figure 1. Pressure-time relationships for the 
Mk 18 grain. 


temperatures. It is apparent that the maximum 
ignition pressure changes only from 550 to 1 ,500 psi 
with a change in propellant temperature from — 24 
to 160 F. The corresponding change in reaction 


52 


IGNITER CONSTRUCTION AND PERFORMANCE 


53 


pressure is from 570 to 2,500 psi. These somewhat 
typical data indicate that ignition pressure is not as 
greatly influenced as reaction pressure by the burn- 
ing rate of the propellant. 

An increase in the quantity of black powder in- 
creases the ignition pressure significantly. Within 
limits, an increase in the relative quantity of black 
powder per unit of free volume in the grain and 
igniter interval decreases the frequency of misfires 
or hangfires at temperatures close to the lower tem- 
perature limit of stable burning for the charge, 
although an increase in the size of the igniter be- 
yond that necessary to produce an ignition pressure 
of approximately 1 ,000 psi does little to decrease the 
temperature at which reliable ignition can be 
obtained. However, an increase in igniter charge 
beyond this point or an increase in propellant tem- 
perature decreases the ignition delay, as is evident 
from Figure 2. 

It should now be emphasized that the ignition 
pressure does not correspond to the pressure ob- 
tained when the igniter is fired in a free space of 
identical geometry involving only inert materials. 



1200 1 600 2000 2400 2800 3200 

PRESSURE IN PSI 

Figure 2. Relation between ignition pressure and 
ignition delay for 2.25-iri. rocket motors. 


In such instances the pressure within the chamber 
rises to perhaps 100 psi because of the reaction of the 
igniter alone. However, in combination with a grain 
of double-base propellant, the ignition pressure may 
be 1,500 psi. These values indicate the effect which 
the presence of the products of reaction of the 


igniter have upon the rate of reaction of the 
ballistite. 


IGNITER CONSTRUCTION AND 
PERFORMANCE 


A typical design of an igniter for a 2.25-in. rocket 
motor 4 is presented as Figure 3. The black powder 
is ignited by an electric squib. Experimental work 



Figure 3. Plastic-case igniter for 2.25-in. rocket 
motor. This unit contains 12 g FFFG black powder. 


has shown that the following is the approximate 
time schedule for the several steps in the ignition 
process. The values given represent the elapsed 
time in milliseconds from the application of the 
electric energy to the squib. 


Melting of bridge wire 3 to 4 

Initiation of black powder 5 to 6 

Rupture of case 18 to 25 

Ignition of propellant charge 25 to 36 


It appears from this time schedule that the 
actual ignition of the propellant charge requires 
approximately one-third of the total ignition period 
and that the remainder is consumed in the action of 
the squib, the initiation of the reaction of the black 
powder, and the rupture of the case. 

Black powder igniters are relatively cheap to 
prepare and involve materials that are readily 
available. In general, either a glazed or shell powder 
of approximately FFF granulation can be employed 
to advantage. It has been found that a decrease 
in the size of the particles to “dust” does not sig- 
nificantly decrease the ignition delay and often re- 
sults in unsatisfactory performance because of the 


54 


IGNITION 


tendency of the dust to cake if slight quantities of 
moisture gain entrance to the igniter. On the other 
hand, efforts to sustain the ignition pressure by the 
use of coarse granulation did not prove particularly 
effective, and it appears that there is little to be 
gained by the use of a granulation coarser than that 
which will permit complete reaction of the black 
powder before expulsion from the rocket motor. 

Since black powder is somewhat hygroscopic, 5 it 
is desirable to seal the igniter case in such a fashion 
as to prevent the entrance of moisture during stor- 
age. Small quantities of water up to approximately 
1.5 weight per cent do not seriously affect the igni- 
tion characteristics (see Figure 4), but an increase 
in the water content of the black powder significantly 
above this value results in erratic and unpredictable 



Figure 4. Influence of moisture in a black powder 
charge upon ignition delay for a 5.0-in. spin-stabi- 
lized rocket. 


ignition delays and may cause disintegration of the 
active ingredients of the squib. A water content of 
1 .5 per cent corresponds to equilibrium at a relative 
humidity of about 92 per cent at 80 F. 

The electric squibs employed in most of the ig- 
niters with which this group has been concerned 
were of a standard deflagrating type prepared by a 
commercial munitions manufacturer. The current 
required was approximately 0.5 ampere in order to 
cause the bridge wire to fail in 3 or 4 milliseconds. 
If currents significantly less than 0.5 ampere were 
used, the time required for the failure of the bridge 
wire was uncertain and increased rapidly until it 
exceeded 1 second with currents of approximately 
0.2 ampere; but increasing the current above 
approximately 1 ampere did not significantly affect 
performance. Squibs can be prepared requiring 


much smaller energies than those indicated above; 
for example, experimental squibs have been tested 
which give reproducible ignition delays with energy 
requirements of less than 20 ergs. 

The squibs employed in many of the rockets 
developed by this group during World War II were 
susceptible to ignition by high-voltage electric dis- 
charge. It was found that the voltage applied be- 
tween the face of the squib and one of the leads 
differed markedly from unit to unit, apparently 
because of irregularity in the depth to which the 
bridge wire was immersed in the active ingredients, 
and that normal statistical variation resulted in a 
limited number of squibs which may have been sen- 
sitive to the static discharges likely to be encoun- 
tered in handling. However, in the course of loading 
several hundred thousand squibs, only two ignitions 
occurred which may be attributed to static discharge . 

The squibs were fired by means of a low-voltage 
electric circuit, part of which was located within the 
rocket motor. The connection between the interior 
of the rocket motor and the leads to the firing circuit 
was accomplished in a number of ways, depending 
upon the design of the particular motor; but the 
maintenance of an adequate seal to prevent the 
entrance of moisture was troublesome. It may 
therefore be desirable in the future to consider the 
use of low-energy squibs and induction firing 6 in 
order to avoid the necessity of sealing the leads and, 
particularly in the case of rockets fired from auto- 
matic launchers, connecting the rounds to the firing 
circuit. The possible hazards arising from stray 
electromagnetic fields may be minimized by the use 
of specially wound coils requiring unusual configu- 
rations of field in order to induce the requisite 
energy in the interior circuit. 

Because of the relatively fragile nature of the 
squib and the black powder grains, it is customary 
to assemble the igniter in some kind of semirigid 
container. From the standpoint of short ignition 
delay it is probably desirable to maintain the ratio 
of the surface of the container to the volume of the 
container as small as possible, but small digressions 
from the spherical shape which is thus indicated do 
not materially influence performance. Igniter cases 
have usually been prepared from plastics 4,7,8 and 
metal. Igniters with tin plate cases 9 have proved 
to be entirely satisfactory with motors having 
nozzles large enough not to be plugged by fragments 
of the case; a typical design for use with a 5.0-in. 
rocket motor is shown in Figure 5. Diffusion of 


SUMMARY OF REQUIREMENTS 


55 


nitroglycerin from the double-base propellant causes 
deterioration of plastic cases; but, under normal 
conditions of storage in Service use, the ballistic 
performance of the igniters does not seem to be 
modified significantly. 

Some type of very thin metal igniter case of 
cylindrical shape would appear to be satisfactory 
for internal-burning grains, and it is possible that 
an alloy of relatively low melting point might be 


greater the stresses imposed upon the grain at the 
time of the rupture of the case; however, up to a 
certain point an increase in the weight of the case 
decreases and renders more reproducible the igni- 
tion delay. Under certain circumstances cloth bag 
igniters appear to deteriorate more rapidly when 
subjected to vibration than do either the metal or 
plastic units. 



Figure 5. General arrangement of Mk 14 igniter 
for a 5.0-in. rocket motor. 


desirable to avoid the difficulties associated with 
nozzle plugging. To facilitate loading and prevent 
movement of the squib with respect to the case 
during vibration, a small stamping or other piece 
should be provided to hold the squib in place. In 
general, the heavier the wall of the igniter case, the 


63 SUMMARY OF REQUIREMENTS 

For optimum performance an igniter should ini- 
tiate the reaction of a propellant charge in a mini- 
mum of time. Apparently the time required to 
initiate the reaction of double-base propellant is 
from 6 to 10 milliseconds, and this time is roughly 
independent of the quantity of black powder em- 
ployed. Normal igniters of the present types 
usually give ignition delays from 25 to 36 milli- 
seconds, depending upon the geometry of the rocket 
motor and the design of the igniter. It is doubtful 
whether the reaction of the propellant charge of a 
rocket motor with a single igniter can be initiated in 
much less than 12 milliseconds; and the decrease of 
ignition delay to this value must be accomplished 
for the most part within the igniter itself. 

An igniter should not react with sufficient violence 
to place undue stresses upon the propellant charge. 
For this reason it is desirable to make the case of the 
igniter no heavier than is necessary to confine the 
ignition charge until it is ignited. Approximately 
6-mil tin plate appears heavy enough to meet this 
requirement. 

An igniter should function over a range of tem- 
peratures which corresponds to the range of suc- 
cessful operation of the round as a whole and should 
also be reasonably resistant to the influx of moisture . 
These, as well as the other requirements summarized 
above, can be satisfied with either plastic case or 
metal case igniters of suitable design. 


Chapter 7 

DRY -PROCESSED DOUBLE-BASE PROPELLANTS 

By B. H. Sage 


71 CLASSES OF PROPELLANTS a 

T he nomenclature associated with the desig- 
nation of the several types of double-base 
propellants is not entirely clear. For present 
purposes they will be considered in two general 
classes: those which are processed by the use of 
solvents, and those which are processed from mix- 
tures with water to the finished propellant without 
the use of solvents. The first class will be referred 
to as solvent-processed propellants and the second 
as dry-processed propellants. Although the cost of 
manufacturing propellants by either of the two 
methods is comparable, the removal of solvent from 
grains having a web thickness greater than 0.5 in. 
requires such unusually long periods of time and 
dimensional uniformity decreases to such an extent 
that grains with thick webs are usually processed by 
the dry method. The dry processing probably in- 
volves a slightly greater hazard during manufacture 
but yields a product of good dimensional uniformity 
which may be prepared in web thicknesses lim- 
ited only by the scale of the available extrusion 
equipment. 

72 COMMENTS ON MANUFACTURING 
METHODS 

No effort will be made in this report to discuss 
the relative merits of the several methods of prepar- 
ing dry-processed propellants, but a few general 
comments appear to be in order. Although con- 
ventional methods are used in the manufacture of 
the requisite nitroglycerin and nitrocellulose, it has 
been found that the nitration and source of the 
cellulose influence significantly the physical char- 
acteristics as well as the ballistic potential of the 
propellant. Double-base powders prepared from 
nitrocellulose made from wood pulp are much more 
difficult to extrude than powders of identical com- 
position prepared from nitrocellulose made from 
cotton linters. For this reason most of the nitro- 

a See also Part III. 


cellulose employed in the manufacture of dry- 
processed double-base propellant in this country 
during World War II was prepared from cotton 
linters. In so far as is known to the writer, the 
reasons underlying this difference in extrusion char- 
acteristics are not yet clear. However, it is evident 
that nitrocellulose with a wood pulp base yields a 
powder which tends to check and crack upon 
extrusion and which gives a much higher velocity 
distribution across the die than powder derived from 
nitrocellulose with a linters base. 

In the case of the slurry process, the nitrocellulose 
is mixed with a relatively large quantity of water 
and agitated. The nitroglycerin is then introduced, 
together with certain of the additive ingredients, 
and the whole permitted to come to substantial 
equilibrium. The nitroglycerin is assimilated by 
the nitrocellulose. The resulting solid or plastic 
phase is separated from the water by means of 
centrifuges. At this point in the process the paste 
contains approximately 30 per cent water by 
weight. It is allowed to age and dry in bags, where 
the moisture content is reduced to approximately 
6 per cent. After blending, the material, which is 
now called “dry paste,” is placed upon differential- 
speed rolls of a design adapted from the rubber 
industry and rolled sufficiently to colloid the stock 
reasonably well. It is then removed from the dif- 
ferential-speed rolls and transferred to even-speed 
rolls, where further mechanical energy is added in 
the course of a number of “bookfolding” operations. 
The resulting sheet is approximately 0.050 in. thick 
and slightly translucent, although the addition of 
approximately 0.2 per cent carbon black renders it 
relatively opaque. 

The details of the manufacturing 15 of double-base 
dry-processed propellant varied significantly from 
plant to plant in accordance with the availability of 
facilities; nevertheless, there appeared to be no 
marked variation in the quality of the product. A 
relatively large number of fires occurred in the 
course of the rolling operation. However, the use of 
special deluge equipment reduced the number of 

b All under Ordnance Department contracts. 


56 


EXTRUSION OF DOUBLE-BASE STOCK 


57 


injuries to personnel to a relatively low value. De- 
tonation of the propellant stock on the rolls has been 
known to occur. 

7 3 EXTRUSION OF DOUBLE BASE 
STOCK 

The extrusion of dry-processed double-base pow- 
der was first carried out in the late fall of 1941. 1 
Additional work was done on a somewhat larger 
scale shortly thereafter, 2 and relatively large grains 


of JP, JPN, and JPH (see Table 2 of Chapter 5) 
sheet stock into finished grains. 

The extrusion operation involves the heating of 
the sheet powder to a temperature of from 100 to 
140 F, depending upon the grain section to be pre- 
pared, and the insertion of the sheet stock as a 
“carpet roll” or as flakes into a horizontal or vertical 
press. After the press has been closed and the pres- 
sure lowered to approximately the vapor pressure of 
water at the charge temperature, the volume of the 
charge is reduced until the charge is extruded at 
pressures from 4,000 to 9,000 psi. 


o 



Figure 1. General arrangement of 18-in. vertical extrusion press at Naval Ordnance Test Station, Inyo- 
kern. 


were extruded at a somewhat later date. 3 A small- 
scale extrusion plant was designed for the Navy 
Department. 4 This was built and operated by the 
Navy at the Naval Powder Factory, Indian Head, 
Maryland. The methods of preparing more complex 
multi web grains 5 are not particularly difficult, and 
there are indications that conventional die design as 
practiced by the plastics industry may be employed 
in the extrusion of a number of the double-base dry- 
processed propellants. A description of the experi- 
mental production facilities developed in the Pasa- 
dena area by CIT under OEMsr-418 is available. 6,7 
These facilities c were used mainly for the processing 

c Most of this equipment has been moved to the Naval 
Ordnance Test Station, Inyokern, California, and to Picatinny 
Arsenal. 


Commercial manufacturers d throughout the coun- 
try utilized horizontal extrusion presses varying in 
diameter from 8 to 15 in. But such presses are dif- 
ficult to feed with other than “carpet roll” extrusion 
charges; hence nearly all material to be reworked in 
the commercial establishments was rerolled into 
sheet stock and in most instances blended with a 
certain amount of new “dry paste.” On the other 
hand, the group at the California Institute has 
generally favored the use of vertical presses because 
they permit the direct extrusion of rework material 
without an intermediate rolling step. Charges of 
double-base stock which had been cut into rela- 
tively small pieces were fed to the vertical presses 


d All under Ordnance Department contracts. 



58 


DRY-PROCESSED DOUBLE-BASE PROPELLANTS 


without difficulty. However, the vertical presses 
required more extensive barricades 8,9 than would 
have been necessary for horizontal presses. It is 
probable that each type of hydraulic press has its 
advantages and shortcomings for a particular 
situation. 

Although it does not appear desirable to enter 
into a detailed discussion of the design and opera- 
tion of extrusion presses, a schematic drawing and a 
photograph of an 18-in. vertical extrusion press are 
presented as Figures 1 and 2. This press, which is 
located at the China Lake Pilot Plant of the Naval 


that the dies be prepared with a relatively high pol- 
ish in order to decrease the friction between the 
metal surface and the propellant being extruded. 
This avoids localized excessive temperatures at 
the interface, which have probably caused at least 
one press ignition. 10 A low coefficient of friction also 
decreases the velocity distribution within the pro- 
pellant during extrusion and consequently may re- 
duce the frequency of inhomogeneities in the 
extruded product. 

The extruded grains immediately undergo a sig- 
nificant change in size, which is usually a dimen- 



Figure 2. Eighteen-inch vertical extrusion press at Naval Ordnance Test Station, Inyokern. 


Ordnance Test Station, Inyokern, California, was 
designed and the equipment constructed by Section 
V of OEMsr-418. The installation requires relatively 
large barricades in order to permit operation with 
propellants of marginal compositions for which the 
frequency of ignition during extrusion would be 
unusually high. 

Double-base dry-processed propellant can be ex- 
truded from dies made up of conical and cylindrical 
sections, or from more complex configurations which 
give lower rates of shear within the propellant for a 
given extrusion velocity. A die used in a 12-in. 
vertical extrusion press 8 for the preparation of the 
Mk 13 grain is illustrated in Figures 3 and 4. 
Since the Mk 13 grain is of cruciform section, no 
“stake” is required; one is necessary, however, with 
a die for an axially perforated grain. It is essential 


sional increase with respect to the die and which in 
some cases may amount to as much as 8 per cent. 
Furthermore, relatively high stresses remain which 
are relieved but slowly at room temperature. There- 
fore, a grain that has not been annealed will gradu- 
ally shorten and become larger in cross section. In 
order to avoid this difficulty, which may exert a 
significant influence on the upper safe-operating 
limit of the charge, the grains are annealed in a low- 
velocity airstream held at 140 F for a period of 
approximately 4 hours per inch of web thickness. 
During this period the grains are placed upon racks 
which assist in eliminating any unusual longitudinal 
curvature. An extruded dry-processed double-base 
propellant grain so supported as not to be deformed 
by its own weight will be straight within approx- 
imately 0.05 in. per ft of length. 


MACHINING 


59 


74 MACHINING 

After being extruded and annealed, the grain is 
subjected to such machining operations as may be 
requisite. In general, it is not necessary to machine 
the periphery of the grain, since in this respect it is 
possible during extrusion to hold the dimensions 
within the limits imposed by ballistic requirements. 


higher tool speeds and feed rates than are employed 
for metals. Some success has been realized in the 
use of plastic saws, but these give a chip with a 
somewhat higher specific surface than is obtained 
by turning or milling. Fires during machining 
operations are relatively rare and can usually be 
traced to foreign material in the ballistite, exceed- 
ingly dull tools, or the inadvertent relative motion 



4 — 2 

HEAD CAP SCREW 


16 
ALLEN 


* DIST 

DIA 

y. DIST 

DIA 

0.000 

2.710 

4.0 0 0 

3.398 

0.400 

2.717 

4.400 

3.546 

0.800 

2.737 

4.800 

3.716 

1.200 

2.772 

5.200 

3.913 

1.600 

2.820 

5.600 

4.144 

2.000 

2.882 

6.000 

4.423 

2.400 

2.957 

6.400 

4.766 

2.800 

3.047 

6.575 


3.200 

3J50 



3.600 

3.267 




GROOVE INCREASED TO CORRECTED 
DEPTH SHOWN AND FAIRED IN BY 
HAND OPERATION AFTER TURNING 
DIE CURVE PROFILE TO DIMENSIONS 
SHOWN IN TABLE 


Figure 3. General arrangement of die used in extrusion of Mk 13 grain. 


However, it is usually necessary to bring the grain 
to a given length and weight within relatively small 
tolerances. Furthermore, it is often desirable to 
apply a plastic support to the end of the grain to aid 
in the distribution of the setback and friction forces 
over its cross section. In order that the cellulose 
acetate or ethyl cellulose reinforcement may be 
bonded satisfactorily, the surfaces of the propellant 
and the plastic must match closely. For this reason 
the grain is usually faced or sawed rather than cut. 

Most double-base propellants can be machined 
readily with conventional machine tools at much 


of metal and propellant surfaces which are in 
contact. 

For the most part, the weight of a propellant 
grain can be held within the desired limits by 
appropriate control of its cross section and length. 
It is usually possible to machine grains to a fixed 
length or into groups of fixed length determined by 
grading the several grains according to their cross 
sections. The specific weight of extruded double- 
base propellant is remarkably constant for a given 
composition; in fact, for JPN powder it is generally 
within 0.5 per cent of 100.5 lb per cu ft. 



60 


DRY-PROCESSED DOUBLE-BASE PROPELLANTS 


75 INHIBITING 

The change in burning area as the reaction pro- 
gresses can be controlled within limits by modifying 
the cross section of the grain. However, in the case 
of a simple external-burning grain such as the cruci- 
form section, which is typified by the Mk 13 and the 
Mk 18 grains, it is not possible to obtain a neutral or 
progressive burning surface without preventing or 



Figure 4A. Die used in extrusion of Mk 13 grain. 

inhibiting burning at certain points on the surface 
of the grain. It has been found that it is a relatively 
simple matter to prevent the surface reaction of 
double-base propellant by the application of suit- 
able coatings. The application of strips or sheets of 
cellulose acetate was the method commonly used in 
this country for inhibiting. 

In the case of the cruciform grain, the arms were 
inhibited on the periphery, as shown in Figure 5 for 
the Mk 13 grain. In general, the thickness of the 
inhibitor was increased with the web thickness and 
was approximately 0.10 in. for the Mk 13 grain. 


The cellulose acetate inhibitor strip was prepared 
by extrusion and applied by the use of solvents 
miscible with both propellant and strip. Numerous 
solvents are suitable; and, as a matter of con- 
venience, mixtures of Cellosolve and methyl Cello- 
solve (2-ethoxy- and 2-methoxy-ethanol) were em- 
ployed. The relative quantities of these compounds 
were varied, depending upon the temperature of 
propellant at the time of application, in order to 
obtain the desired rate of softening of the cellulose 
acetate. The strips were applied manually to most 
of the grains used in Service rockets. Efforts have 
been made to develop automatic equipment for this 
purpose, since undesirable physiological effects are 
usually experienced by operating personnel as a 
result of either the solvents themselves or the nitro- 
glycerin in the propellant. However, such equip- 
ment has not yet proved entirely satisfactory. 

The British used a slightly softer material con- 
sisting of cellulose acetate and triacetin, i.e., dummy 
cordite. It is believed that the development of a 
suitable plastic of very low elastic limit might permit 
the use of automatic machines for the application 
of the inhibitor strips and still avoid the difficulties 
which otherwise result from lack of uniformity in 
the curvature of the grains and strips. However, 
the inhibitor strip on the completed grain must be 
hard enough not to be unduly deformed during 
normal handling and storage. 

In the case of end-burning or internal-burning 
grains where the entire periphery is inhibited, two 
techniques appear to be promising. One involves 
wrapping the grain with relatively thin cellulose 
acetate, or perhaps other plastic sheet, using ap- 
propriate plasticizers to obtain a satisfactory bond 
between the propellant and the sheet. This method 
may be modified to permit thin plastic tape to be 
wrapped spirally on the grain. Experience with the 
latter form of inhibitor indicates that larger geo- 
metric irregularities may be permitted than can be 
tolerated with full-width sheet. The second ap- 
proach involves extruded or molded tubing which is 
shrunk or molded onto the grain; but, in so far as the 
writer is aware, these techniques, which were de- 
veloped by the British, have not been widely used 
in this country. e The end of an inhibited internal- 
burning cylindrical grain is shown in Figure 8 of 
Chapter 5. 

e The experience of the Allegany Ballistics Laboratory with 
these methods of inhibiting is indicated in Section 11.2.1 of 
this volume. 


STABILITY OF BURNING 


61 


76 CHEMICAL STABILITY 

The stability of propellants is of importance in 
connection with their storage. Of the principal com- 
ponents, it is probably the nitrocellulose which con- 
tributes most to the chemical instability of multi- 
base propellants. Stabilizers are therefore employed 
to avoid the accumulation of the oxides of nitrogen 
which results from the spontaneous decomposition 
of the nitrocellulose, since the presence of these free 
oxides of nitrogen apparently accelerates further 
decomposition. 


combination of nitrocellulose and stabilizer which 
results in a propellant with a long storage life under 
adverse conditions, rather than the choice of a par- 
ticular stabilizer. 

Thick-webbed grains of dry-extruded double- 
base propellant involve a somewhat unique problem 
with regard to stability, since it appears that small 
quantities of the gases from the decomposition of 
the nitrocellulose do not react completely with the 
stabilizer. In the case of thin-webbed grain, these 
gases diffuse to the exterior surface and cause no 
particular difficulty. With thick- webbed grains, 



Figure 4B. Die used in extrusion of Mk 13 grain. 


In JP propellant, which was developed primarily 
for use with trench mortars, the stabilizer was 
diphenylamine. However, it is now evident that 
this material is unduly active and tends to acceler- 
ate the decomposition of the nitrocellulose. Of the 
relatively large number of stabilizers that are avail- 
able, ethyl centralite seems to be the optimum for 
the stabilization of colloidal propellants with a nitro- 
glycerin-nitrocellulose base. This material can be 
incorporated readily into the propellant, and the 
equilibrium at constant pressure is such that the 
oxides of nitrogen do not contribute to any sig- 
nificant extent to the decomposition of the nitro- 
cellulose. It is probably desirable to continue the 
investigation of stabilizers; but it is believed that 
studies of the characteristics of nitrocellulose will 
contribute equally, if not more, to the stability of 
double-base propellants, for it is apparently the 


however, the fugacity of this material within the 
grain may reach relatively large values. The cor- 
responding mechanical stresses result in cracking of 
the grain, usually parallel to the axis, or spalling on 
the surface. Since the occurrence of such defects 
during storage may well become a factor limiting 
the size of large propellant grains, it is believed 
that improvement in the effectiveness of stabilizers 
will result in an increase in the size of grain which 
it will be feasible to manufacture and store. 

77 STABILITY OF BURNING 

In the study of the deflagrating characteristics of 
propellants it has been found that irregular reaction 
pressures are often encountered, especially with pro- 
pellant of relatively high burning rate, such as JPN 
or JPH. When efforts were made to utilize charges 


62 


DRY-PROCESSED DOUBLE-BASE PROPELLANTS 


of these materials which were permitted to burn 
from extended plane surfaces, the reaction was suf- 
ficiently erratic to cause mechanical failure of the 
grain. For this reason it was impossible to use single 
continuous strips of inhibitor on the periphery of 
the arms of the Mk 13 and Mk 18 grains. More- 
over, attempts to obtain regular burning within a 
cylindrical annular perforation also resulted in 
mechanical failure of the grain because of the in- 
stability of the reaction. This situation was over- 
come by drilling radial holes at somewhat random 
intervals but spaced longitudinally not more than 
1 in. apart. It has recently been found that rela- 


sufficient background of empirical information is 
available, however, to permit the design of pro- 
pellant charges of each of the powders commonly 
employed. 

7 8 SUMMARY AND RECOMMENDATIONS f 

The general status of the knowledge relating to 
dry-processed double-base colloidal propellants that 
have been employed in artillery rockets has been 
indicated in the foregoing discussion. It now ap- 
pears that rocket motors having an overall specific 
impulse of approximately 110 lb-sec per lb can be 



LINEAR DIMENSIONS 
ARE IN INCHES 


m ^94 m 9.00 

33.75 AVERAGE ► 


^ ^ 7*88 94 -* 

T77777X 

K/, INHIBITOR '//I Y///Z 

*• 7.88 *+•" 

9.00 7.88 H 

ZZZZZZZZZLL 

v// /// /SA ~~ 

- i— Ck r\ r\ - ■ - 7 QQ y 

4 .DO m \ m i.oo -y— ..w 

y.uu 

tV/'/V/VZ i 'VZZZZZZZZ. 



DEVELOPED SURFACE AT 2"73 DIA 


Figure 5. General arrangement of inhibitors on Mk 13 grain. 


tively stable burning occurs in a star-shaped 
perforation. 

With propellants of intermediate burning rate 
(0.4 ips at 70 F and 1,000 psi), of which British 
cordite is typical, it is possible to obtain stable 
burning with less frequent interruption of plane 
surfaces, or surfaces of relatively large radius of 
curvature, than is necessary with propellants hav- 
ing high burning rates (0.65 ips at 70 F and 1,000 
psi, for example). In the case of propellants with 
burning rates of less than approximately 0.25 ips at 
70 F and 1,000 psi it has been found that stable 
burning may be obtained with almost any shape of 
grain that does not involve excessive energy ex- 
changes associated with friction. (Instability at- 
tributable to frictional effects is entirely separate 
from the type under discussion.) 

The mechanism of unstable burning not directly 
associated with frictional effects is not thoroughly 
understood, but may be related to resonance. A 


developed which will operate at temperatures be- 
tween — 30 and 130 F. The burnt velocity obtain- 
able with such a rocket motor depends almost 
entirely upon the payload to be carried. Five-inch 
rounds of reasonable length-to-caliber ratio can be 
made with burnt velocities greater than 3,500 fps 
and with payloads of approximately 10 lb. How- 
ever, if the payload is increased until it is equivalent 
to a shell of comparable caliber, velocities in excess 
of 2,500 fps are unlikely. The use of internal- 
burning grains prepared from existing propellants 
seems feasible and not too costly. Such grains per- 
mit rates of spin in excess of 400 rps to be obtained 
at temperatures up to 120 F without failure of the 
grain. 

Regarding propellants with potentials in excess 
of 200 lb-sec per lb, there is little to be gained at 
present by modifying the composition greatly from 


f See Chapter 13 for additional recommendations. 


SUMMARY AND RECOMMENDATIONS 


63 


that of JPN powder. Such propellants apparently 
have insufficient latitude to permit the addition of 
buffer components which will decrease the influence 
of temperature and pressure on reaction rate. How- 
ever, propellants having potentials of the order of 
150 lb-sec per lb are promising in this respect; and 
it is probable that hydrocellulose and magnesium 
oxide in conjunction with potassium nitrate will 
prove particularly useful. It is believed that, for 
the time being, developments requiring propellants 
of intermediate potential may proceed satisfactorily 
on the basis of material approximating the H-4 
composition recorded in Table 2 of Chapter 5. It 
does not appear that any new propellant which 
would justify delaying the program will be available 
within the next year (1947) in sufficient quantities 
for experimental production. Accordingly, it is rec- 
ommended that the development of rocket ord- 
nance involving dry-processed double-base colloidal 
propellants utilize the existing JPN formulation for 
a high-potential, fast-burning powder and the H-4 
formulation which can be dry-extruded for a powder 
of intermediate potential. 

Two lines of endeavor should probably be fol- 
lowed in the further development of dry-processed 
colloidal propellants. In the first place, a careful 
investigation should be made of the so-called buffer 
constituents which appear by their control of the 
chemical equilibrium to decrease the influence of 
temperature and pressure on the burning rate. 
Particular emphasis should be given the application 
of these constituents to the propellants of higher 
potential, with which they do not now appear to be 
sufficiently effective to warrant their use. Such 


studies, together with investigations of stabilizers 
and the character of the nitrocellulose, can well be 
carried out at academic institutions or government 
laboratories. The second approach should involve a 
systematic study of the influence of composition 
upon the physical, chemical, and ballistic charac- 
teristics of a number of systems comprising the 
principal components of existing double-base pro- 
pellants. In this connection it is believed that 
investigation of such restricted ternary systems as 
the nitrocellulose-nitroglycerin-ethyl centralite sys- 
tem and the ethylene glycol dinitrate-nitrocellulose- 
ethyl centralite system is worth while. 

The foregoing suggestions are not intended to 
cover other than the immediate problems of interest 
in the study of dry-processed double-base colloidal 
propellants. There is a large field of research to be 
investigated in the development of new types of 
smokeless propellants that show relatively small in- 
fluences of pressure, temperature, and transfer of 
radiant energy upon burning rate. Furthermore, 
there is the field of liquid propellants which, in the 
opinion of the writer, will probably supplant solid 
propellants in nearly all large rocket-propelled de- 
vices. The caliber of the rocket for which transfer 
from solid to liquid fuels will prove advantageous 
has yet to be established, but it is probable that 
there will be a range of sizes in which the applica- 
tion will determine the choice of a solid- or a liquid- 
fueled device. It is hoped that an effort will be made 
to standardize simple artillery rocket motors in 
order that a relatively wide variety of heads and 
stabilizing equipment can be used with a given 
motor. 






























































• • 




































































































PART III 


ROCKET ORDNANCE: THERMODYNAMICS AND RELATED PROBLEMS 


By R. E. 


P art iii of this volume will be concerned prin- 
cipally with problems arising in the development 
of colloidal solid rocket propellants and is really a 
summary of many of the final reports issued from 
Allegany Ballistics Laboratory [ABL], which was 
operated by George Washington University under 
contract 15 with the Office of Scientific Research and 
Development with technical supervision by Section 
H, Division 3, NDRC. Much of the pioneering 
work was done by the Section H group working at 
the Naval Powder Factory, Indian Head, Mary- 
land, from 1941 through 1943. In this phase of the 
work close cooperation was established with the 
Hercules Powder Company, which, under contract 
first with OSRD and later with the Ordnance De- 
partment, contributed greatly to the phases of the 
program lying between development and produc- 
tion. Laboratory experimental work and theoretical 
studies on propellants were carried on by groups at 
the Bell Telephone Laboratories, University of 
Minnesota, University of Wisconsin, and Duke 
University, which worked very closely with the 
central Section H Laboratory, first at Indian Head, 
afterwards at Allegany. All these agencies con- 
tributed to the developments described in the fol- 
lowing. Notable contributions to the general sub- 
ject were made by Section L, Division 3, NDRC, 
Division 8, NDRC, Division 1, NDRC, the Bureau 
of Ordnance, U. S. Navy, and the Rocket Develop- 
ment Division, Ordnance Department, U. S. Army. 
These are discussed systematically elsewhere and 
will only be referred to casually in this report. 

The problems in the physical chemistry of rocket 
propellants discussed in this report all arose from 
very practical questions which had to be solved in 
the development of rockets. These problems fall 
into categories well known in physical chemistry, 
namely, thermodynamic problems, kinetic prob- 
lems, and structural problems. In Chapters 9, 10, 


a Director of Research, Allegany Ballistics Laboratory. 
b Contract OEMsr-273. 


Gibson * 


and 11 the studies of rocket propellants will be 
summarized under each of these headings, respec- 
tively, and in each chapter an attempt will be made 
to indicate, first, the practical problems in the 
functioning of rockets that were encountered, sec- 
ond, the problems in the physical chemistry of the 
propellants that arose from these functional prob- 
lems, and, third, a summary of the results obtained. 
Chapter 12 will give a short summary of the applica- 
tion of these problems to internal ballistics. Chap- 
ter 13 will give a summary of the rocket propellants 
which were developed by V-J Day and will indicate 
lines along which progress will probably be made 
in the future. As far as possible, reference to the 
original detailed reports will be given. 

The reader who is unfamiliar with rocket prob- 
lems is urged to consult Rocket Fundamentals , l a 
composite report to which a number of development 
agencies contributed and in which a fairly complete 
but elementary exposition of the principles of rocket 
design and action is given. 

Since this volume will be printed long after 
V-J Day, it is fitting to point out that a great deal 
of the work described here has been continued with 
excellent results since the NDRC activities stopped. 
Allegany Ballistics Laboratory has continued opera- 
tions with a new contractor, the Hercules Powder 
Company, under contract with the Bureau of Ord- 
nance, U. S. Navy. The results and techniques 
developed at the laboratory under NDRC have 
been applied and extended to the development of 
large rockets with solventless-extruded and cast 
double-base powder charges, and devices of great 
interest in the guided missiles program are being 
perfected. Reports from this laboratory should be 
consulted for the sequel to this summary. Further- 
more, the laboratory and theoretical studies of pro- 
pellants conducted at the University of Minnesota 
have continued under the auspices of the U. S. 
Navy. Reports from this university should also be 
consulted for the continuation of the work started 
by NDRC. 


65 































•• 




































on 







Chapter 8 

TYPES OF ROCKET PROPELLANTS 

By R. E. Gibson 


81 JET PROPULSION, ROCKETS, AND 
PROPELLANTS 

I n modern civilian or military engineering there 
is a wide variety of devices for propelling projec- 
tiles or other vehicles. Although they may differ 
widely in their construction and in other superficial 
respects, practically all propulsion mechanisms 
have fundamentally the same basis: they depend 
on the conversion of the energy of a controllable 
chemical reaction into elastic energy of a gas, which 
is then converted, by a suitable mechanical device, 
into kinetic energy of motion in a given direction. 
The mechanical devices by which the elastic energy 
of the gases is converted into useful work, i.e., 
into the kinetic energy of the vehicles, vary in com- 
plexity from the locomotive or airplane engine to 
the simple gun barrel or rocket jet. The choice of 
engine depends mostly on the ultimate application; 
the rate at which energy must be supplied, the 
mobility of the apparatus, the number of hours of 
working life required, and other performance re- 
quirements must be balanced against economic 
factors in choosing the chemical reactants and the 
mechanical apparatus to be used . Few people would 
use nitrocellulose powder to fire a locomotive, and 
few would use coal to propel a large military missile. 


Rockets 

Probably the simplest device for converting the 
clastic energy of a gas into the directed kinetic 
energy of a vehicle is the jet engine. Like other 
motors, these jet engines depend for their energy on 
a chemical reaction which we may consider to be an 
oxidation reaction involving a fuel (the substance 
to be oxidized) and an oxidizing agent. 

A rocket is a jet-propelled vehicle which carries 
with it all the components needed for the energy 
producing chemical reactions, i.e., both the fuel 
and the oxidizer. This characteristic differentiates 
the rocket from other jet engines such as the ram 
jet, the pulse (or reso) jet or the turbo jet, all of 


which draw their oxidizer from the atmosphere 
through which they pass. The ram jet and pulse jet 
draw in air simply by making use of the dynamic 
pressure produced by their motion through the air, 
whereas the turbo jet makes use of compressors 
driven by part of the energy generated by the 
motor. 

82 ROCKET PROPELLANTS 

The term rocket propellant is applied to the 
chemical substance or substances which react to 
produce the hot gases whose elastic energy is to be 
converted into the kinetic energy of motion of the 
projectile. There are two main types of rocket pro- 
pellants: liquid propellants and solid propellants. 

821 Liquid Propellants 

Liquid propellants in turn fall into two main 
classes: bi-fluid systems and mono-fluid systems. In 
bi-fluid systems, which have found most common 
use to date, the oxidizer and the fuel are kept in 
separate tanks in the rocket and fed in proper pro- 
portions into a combustion chamber where they 
react. Such systems are relatively safe as regards 
hazards during storage or transit and permit a wide 
range of control of rate of gas evolution and tem- 
perature, because it is possible to control inde- 
pendently the supply of fuel and oxidizer. Typical 
oxidizers are nitric acid, hydrogen peroxide, and 
liquid oxygen, and typical fuels are aniline (or 
mixed aromatic amines) , hydrazine, methyl alcohol, 
and gasoline. Mono-fluid rocket propellants are 
liquids which contain in themselves sufficient oxy- 
gen to give fairly complete oxidation of the other 
elements with evolution of heat, when a reaction is 
started. Although all such substances are of neces- 
sity thermodynamically unstable, a number of suit- 
able propellants, such as nitro-me thane and hydro- 
gen peroxide, have been found which decompose at 
a negligible rate at ordinary temperatures and can, 


67 


68 


TYPES OF ROCKET PROPELLANTS 


therefore, be handled with comparative safety. 
Nevertheless, precautions required to store and 
handle such propellants tend to offset the obvious 
engineering advantages to be gained when two 
liquids are replaced by one. 

82,2 Solid Propellants 

In solid propellants the fuel and the oxidizer are 
intimately mixed and in a condition to react rapidly, 
but controllably, when the necessary activation 
energy is supplied, usually by a device called an 
igniter. It is a necessary characteristic of all solid 
rocket propellants that the reaction (which is usually 
called the “burning”) take place only on the ex- 
posed surfaces of the solids and that burning 
proceed in directions normal to the surfaces at a 
rate which is the same at all points. 

For very large and long-range rockets such as the 
V— 2, or for applications where good thrust control 
is required, as in a jet plane, liquid propellants 
possess overwhelming advantages over solid pro- 
pellants. It must be noted, however, that the 
valves and plumbing systems in these large rockets 
are complicated and costly; in the V— 2 rocket the 
fuel system must be capable of supplying about 270 
lb of fuel and oxidizer per second. In smaller 
rockets, therefore, particularly where ease of han- 
dling and simplicity of design are important, solid 
rocket propellants have a field of application in 
which they are unrivaled. 

During World War II the activities of Section H, 
Division 3, NDRC, were confined to rockets or jet- 
propelled devices weighing less than 200 lb. Its 
attention was, therefore, concentrated on solid 
propellants, and Part III of this report will be con- 
cerned only with this type of propellant. 

Composite and Colloidal 
Propellants 

Two main classes of solid propellants are recog- 
nized. In one class the oxidizer and the fuel are 
present as separate molecules, or as small crystalline 
aggregates intimately mixed and held together by 
adhesives designed to give suitable mechanical 
properties to the mass as a whole. These are called 
composite propellants, and the classical example is 
ordinary black powder where the oxidizer is potas- 


sium nitrate and the fuel is charcoal. During World 
War II considerable effort was expended in the 
development of new and improved composite pro- 
pellants. Section H, Division 3, took no part in the 
actual development of these propellants but was 
active in testing them ballistically. The research 
and development work was done by Division 8, 
NDRC, and by the Guggenheim Aeronautical 
Laboratory, California Institute of Technology 
[GALCIT]. GALCIT developments were later ex- 
tended and applied by the Aerojet Engineering 
Corporation. Three significant varieties of com- 
posite propellants were developed by these agencies. 
Division 8 produced composite propellants by the 
molding, solvent extrusion, and casting methods. 
GALCIT produced a number of cast perchlorate 
propellants. The preparation and properties of 
these propellants are given in Chapter 13. In 
smaller artillery rockets, composite propellants 
found relatively limited application, although the 
solvent-extruded composites gave the answer to a 
very urgent need that arose in connection with the 
infantry bazooka rocket. 1 On the other hand, the 
composite propellants, because of simplicity of 
manufacture and their desirable burning properties, 
proved to be extremely well suited to use in rocket 
motors where long burning times and large amounts 
of propellant were required. Indeed their only dis- 
advantage arose from the smoke they produced. 

The second class of propellants, and the class 
which found most extensive use in the artillery 
rockets of all nations engaged in World War II, 
comprises the colloidal propellants which have been 
used for years. In colloidal propellants, the oxidizer 
and the fuel are on the same molecule, and the solid 
itself is macroscopically homogeneous. Colloidal 
propellants consist essentially of a high polymer 
which is rich in oxygen and can undergo an exother- 
mic reaction in which its elements are raised to a 
higher state of oxidation. The high polymer may be 
plasticized with oxygen-rich plasticizers which are 
metastable chemically, or with plasticizers which 
are essentially fuels. The plastic formed by the 
interaction of high polymer and the plasticizers 
gives a homogeneous mass in which suitable 
physical properties may be developed. Since the 
main stimulus for improving solid propellants for 
rockets came from the desire to throw heavier pay- 
loads faster and farther for military purposes, it is 
not at all surprising that rocket development 
agencies in all countries should have turned to con- 


ROCKET PROPELLANTS 


69 


ventional gun propellants for the first source of 
high-energy fuels. Of the various gun propellants 
available, the class called double-base powders 
proved most suitable, chiefly because of their 
ability to react reliably at relatively low pressures, 
300 to 1,500 psi, and because it was found possible 
to fabricate them into “grains” of suitable shapes 
and sizes for rocket work. Single-base powders 
possess neither of these properties. 

Double-base powders receive their name from the 
fact that they contain two explosive ingredients — 
one being a high polymer (up to now always nitro- 
cellulose) and the other being a plasticizer, usually 
nitroglycerin; other explosive plasticizers have also 
been used, e.g., diethylene glycol dinitrate, DINA, 
and TNT. Generally speaking, the nitroglycerin 
forms between 30 and 45 per cent of the whole mass, 
the rest being nitrocellulose with varying amounts 
of auxiliary plasticizers such as ethyl or methyl 
centralite, triacetin, and clinitrotoluene, stabilizers 
such as ethyl centralite or diphenylamine, and inor- 
ganic salts such as potassium nitrate or potassium 
sulphate. In some very desirable double-base rocket 
propellants developed during World War II, the 
amounts of auxiliary plasticizers such as triacetin or 
centralite rose in amount to something between 5 
and 20 per cent of the whole composition. In double- 
base powders the nitrocellulose is gelatinized with 
or without the help of an active volatile solvent by 
mechanical working. The resulting mass is a hard, 
hornlike, homogeneous, rigid colloid which obeys 
ideally the law of burning in parallel layers. 

For rocket applications where short burning times 
and high accelerations are required, double-base 
powder gelatinized with the help of an active vola- 
tile solvent is very suitable because of the high 
physical strength that may be developed in the 
grains. The “solvent process,” although also being 
advantageous because of the ease and relative safety 
in manufacture, is severely limited in application, 
since the removal of the solvent sets an upper limit 
to the “web” thickness (minimum dimension of 


grain) that may be obtained. In the “solventless 
process” the double-base powder is gelatinized by 
severe working on heated rolls without the aid of an 
active volatile solvent. This method is particularly 
advantageous when longer burning times are re- 
quired. In the solventless process the colloided 
powder is formed into grains by extrusion under 
high pressure at elevated temperatures, and essen- 
tially the only upper limit to the web thickness that 
can be made available is that imposed by the size of 
press that is safe and practical to operate. The 
process is not suitable for making single-base powder 
but is well suited to the manufacture of double-base 
powder containing less than 60 per cent nitro- 
cellulose. 

The solventless process was developed in Germany 
prior to World War I and was introduced into 
Great Britain and France shortly thereafter. It was 
extensively used in Russia at least as early as 1931 . 
Prior to World War II only a small amount of 
solventless double-base powder was used in the 
United States, and this only in sheet form for use in 
trench mortars. No apparatus existed for extruding 
solventless powder into cylindrical grains, and in- 
deed the industry exhibited a strong prejudice 
against setting up such an operation. Thus, while 
the rocket developers in Great Britain found in 1935 
a ready production source of a high-power solid 
rocket propellant in the factories used for making 
solventless cordite for the Royal Navy, the Amer- 
ican rocket developer found himself starting from 
scratch, or rather several yards behind the line. It 
is not too much to say that the setting up of a 
solventless powder industry came directly as a con- 
sequence of the visits of NDRC investigators to 
Great Britain. 

This introductory chapter concludes with a chart 
illustrating the various levels of problems connected 
with the development of a complete artillery rocket. 
It is designed to give the reader a general idea of the 
problems encountered and the equipment and fa- 
cilities needed for their solution. 


70 


TYPES OF ROCKET PROPELLANTS 


Table 1 . Research and facilities required in the development of a rocket motor. 



Research on Metal Parts 

Problems 

New materials for rockets 
Application of new engineer- 
ing techniques to rocket 
manufacture 

Protection of metal parts, 
e.g., thermal insulation 
Control of nozzle erosion 

Facilities 

Engineers in contact with 
engineering companies to 
develop new materials 
Tests of products under fir- 
ing conditions 

Facilities for heat treating 
steel 

Facilities for handling ma- 
terials other than steel 


Research on Propellants 

Problems 

Search for new propellants 
of suitable burning & me- 
chanical properties 

Investigation of new charge 
shapes & means of prep- 
aration 

Study of the mechanism of 
burning of rocket propel- 
lants & the effect of en- 
vironment on the burning 

Study of the flow & thermo- 
dynamic properties of pro- 
pellant gases 

Study of new methods of 
fabrication of propellant 
charges 

Facilities 

Source of powders of con- 
trolled compositions 

Static firing ranges for 
studying burning proper- 
ties under different con- 
ditions of pressure, tem- 
perature, gas flow, & en- 
vironment 

Lab for study of the chemi- 
cal, physical, & mechani- 
cal properties of propel- 
lants 

Facilities for preparing pro- 
pellant charges in a wide 
variety of shapes & sizes 

Theoretical group for de- 
veloping & understand- 
ing internal ballistics of 
rockets 




Chapter 9 

THERMODYNAMIC PROBLEMS 


By F. T. McClure a 


91 ROCKET ACTION— THRUST — 
SPECIFIC IMPULSE OR 
EFFECTIVE GAS VELOCITY 

S ufficiently fine analysis of any propulsion 
system will resolve it into an example of New- 
ton’s third law of motion, namely, “To any action 
there is an equal and opposite reaction.” Rocket 
propulsion is a particularly simple and direct prac- 
tical example of this law. The rocket chamber or 
motor exerts a force on the gases contained therein, 
causing them to be expelled to the rear. This, if one 
wishes, is the action. In turn, the gases exert an 
equal force (in the opposite direction) on the rocket, 
causing it to be propelled forward. This, then, is 
the reaction. 

One may guess (and in fact it is a consequence of 
Newton’s second law of motion) that, for rocket 
motors of the same configuration operating under 
the same pressure conditions and using the same 
fuel, the thrust ( F ) will be proportional to the 
mass rate of exhaust of fuel. The proportionality 
constant is generally called the specific impulse (I) 
or the effective gas velocity {V E ) depending on the 
units in which it is expressed, so that 

F = ml 

or (1) 

F = mV e 

where m is the mass rate of discharge of fuel. In 
this country, I is usually expressed as the pounds 
force for each pound per second mass rate of dis- 
charge, while V E is expressed in feet per second. 
Then 

V E = 32.167. 

Because of the discharge of the fuel, the mass of a 
rocket decreases during the acceleration or burning 
time. If W is the mass of the rocket then dW/dt = 
— m, and according to Newton’s second law of 
motion 

a Former Chief of the Ballistics Design Section of the 
Allegany Ballistics Laboratory. 


F -+v—-vJ% -wf, ( 2) 

where V is the velocity of the rocket. Integration of 
the last of equations (2) from the point of initiation 
of the thrust to the point at which the fuel is con- 
sumed leads to a final velocity b given by 

F 0 = V E In (l + (3) 

where m 0 is the original mass of fuel and M is the 
mass of the rocket without fuel. 

Equation (3) clearly exposes the significance of 
the effective gas velocity or specific impulse to 
rocketry. Obviously, fuels capable of producing 
high specific impulse are most desirable, particu- 
larly for very high-velocity or long-range rockets. 
As discussed in the next section, the specific impulse 
of a fuel is determined partly by the operating con- 
ditions (pressure), partly by the motor geometry 
(expansion ratio), and largely by the thermody- 
namic properties (heat capacities, molecular weight, 
and temperature) of the propellant gas which it 
generates. It is through the specific impulse, then, 
that the thermodynamic properties of a fuel provide 
a measure of its potential. 

It should be emphasized that the thermodynamic 
properties of a fuel are not the only properties of 
significance in determining its desirability. This 
may be seen by further examination of equation (3) . 
A fuel must be packaged, and the weight of the 
package or container is included in M. The greater 
the weight of container necessary for a given weight 
of fuel, the more difficult it is to achieve a high ratio, 
mo/M , of fuel weight to empty rocket weight. The 
container weight, however, is largely determined by 
the volume, and thus a high-density fuel has the 
advantage of a lower ratio of container weight to 
fuel weight. For example, the high specific impulse 
of the liquid hydrogen-liquid oxygen combination 
(due to the low molecular weight of the gases gener- 

b These equations neglect the effect of gravity and air re- 
sistance, both of which must be considered in dealing with 
high-velocity, long-range rockets. 


71 


72 


THERMODYNAMIC PROBLEMS 


ated) is partially nullified by the large tanks re- 
quired for the hydrogen (because of its very low 
density). Such considerations apply to both solid 
and liquid fuels. 

Other properties of fuels are also of importance 
in determining their usefulness. In particular, the 
ease, rapidity, and uniformity with which the com- 
plete conversion of the fuel into the propellant gas 
can be accomplished is important in determining 
the weight of the combustion chamber, which is also 
part of M in equation (3) . This is again true of both 
liquid and solid fuel but is more strongly felt in the 
latter case because here the combustion chamber is 
also the container for the fuel. Such problems fall 
in the field of “interior ballistics.” 

It may also be worth noting that rockets use 
large quantities of fuel so that the ease, cost, and 
hazard associated with the manufacture, storage, 
transportation, and handling are important con- 
siderations in choosing a fuel. 

92 THE CALCULATION OF THE 
SPECIFIC IMPULSE— THE REDUCED 
SPECIFIC IMPULSE 

More careful study of the flow of gas from a rocket 
motor not only verifies the assumptions of the pre- 
ceding section but also elucidates the dependence 
of the specific impulse on the thermodynamic pro- 
perties of the propellant gas, the nozzle geometry, 
and the operating conditions. Such a detailed 
analysis is carried out in Chapter II and Appendices 
2 through 8 of reference 1 . An important result is 
that the specific impulse of a fuel-motor combina- 
tion can be separated into a product of \^nRT c and 
a function of y, P a /P c , and A e /A t . Here, n is the 
inverse of the molecular weight, T c is the absolute 
temperature, y is the ratio of the heat capacities 
at constant pressure and constant volume, and P c 
is the pressure of the gases in the combustion 
chamber, whereas P a is the pressure of the surround- 
ing atmosphere, A e is the area of the nozzle exit, 
A t is the area of the nozzle throat (narrowest sec- 
tion), and R is the universal gas constant. Because 
of this separability the quantity 

I 

'SnRT'c 

which is called the reduced specific impulse , is inde- 
pendent of n and T c , and therefore may be tabulated 


or graphed as a function of the pressure ratio 
(Pa/Pc), the expansion ratio (A e / A t ), and 7 without 
reference to the molecular weight or temperature of 
the gas. In reference 2 the reduced specific impulse 
is tabulated and graphed over a wide range of values 
as a function of the pressure ratio and expansion 
ratio for each of the values 1.15, 1.20, 1.25, 1.30, 
1.35, and 1.40 for 7. The graphs for 7 = 1.20 are 
reproduced in Figure 1 as a sample. This report 2 
also includes sample calculations and a summary 
of formulas with provisions already made for appro- 
priate units, so that it becomes a simple matter to 
estimate the specific impulse for a given pressure 
ratio and expansion ratio providing n, Tc, and 7 for 
the fuel are known. 

A point of caution must be emphasized here. As 
Figure 1 indicates, the specific impulse increases 
with increasing expansion ratio until it reaches a 
maximum. This maximum occurs at the point where 
the exit pressure is just equal to the pressure of the 
surrounding atmosphere. It must not be concluded, 
however, that a large expansion ratio can be ob- 
tained by “opening up” the nozzle rather than by 
increasing its length, thus avoiding a penalty in 
nozzle weight. The graphs given neglect the side- 
ways motion of the gas in the nozzle, which contrib- 
utes nothing to the thrust. This effect is only neg- 
ligible providing the divergence of nozzle is not too 
great . 


93 THE DISCHARGE COEFFICIENT 


As shown in reference 1, the mass rate of dis- 
charge of gas from the rocket motor may be ex- 
pressed in the form 

m = C D A t P c , (4) 


where C D , the discharge coefficient, is given by 


Cd 



(7 + D/[2(y - 1)1 


V7 

VnRTc 


(5) 


and thus is determined by the thermodynamic 
properties of the gas in the rocket chamber. Sample 
calculations in appropriate units are given in ref- 
erences 1 and 2. 

According to equation (5) the discharge coefficient 
is independent of geometry of the rocket motor. 
Actually there is a slight dependence on the geom- 
etry through the ratio of throat area to the free 
area of chamber (i.e., the cross-sectional area not 


THE DISCHARGE COEFFICIENT 


73 


/ 

\ 




Figure 1. Properties of a propellant gas with y = 1.20. A. Reduced specific impulse vs expansion ratio at 
various pressure ratios. B. Reduced specific impulse vs pressure ratio for various expansion ratios. 



74 


THERMODYNAMIC PROBLEMS 


occupied by propellant, etc.,— referred to as the 
“port area”). This dependence is due to the pres- 
sure drop and velocity gradient in the combustion 
chamber. The effect is discussed in some detail in 
Appendix 6 of reference 1 . 


THE THRUST COEFFICIENT 


Frequently it is advantageous to express the 
thrust in terms of the chamber pressure according 
to the equation 

F = C F A t P c , (6) 

where C F is known as the thrust coefficient. By use 
of equations (1) and (4) one may obtain 


so that 

Further, 

obtains 

C F — 


F = C D IA t P c 

C F = C D I. (7) 

combining equations (5) and (7), one 

/ 9 X^+l)/[2(7-D] 

\Y + 1/ Vy (vnRTj (8) 


It will be noticed that the last factor on the right of 
equation (8) is just the reduced specific impulse, 
which is a function of y, P a /P c , and A e /A t . Thus 
the thrust coefficient is a function of y, P a /P c , and 
A e l A i and is independent of the molecular weight 
and temperature of the propellant gas. 

The graphs of reduced specific impulse in refer- 
ence 2 may be used to compute the thrust coefficient 
through equation (8) . A sample calculation is given 
in the reference. 


95 CALCULATION OF THE 

THERMODYNAMIC PROPERTIES OF 
THE GAS FROM THE COMPOSITION 
OF THE FUEL 

From the preceding sections it is clear that the 
important properties of the propellant gas, from 
the standpoint of specific impulse, discharge coeffi- 
cient, etc., are the values of y, n, and T c . Ideally, 
the temperature of gases in the combustion chamber 
is the so-called isobaric adiabatic flame temperature, 
which is related in a simple manner to the higher 
isochoric adiabatic flame temperature characteristic 
of the reaction in a closed vessel. Both adiabatic 


flame temperatures are, in theory, calculable from 
the thermodynamic properties of the fuel. Their 
definitions and relationship are discussed in Appen- 
dix 2 of reference 1 . 

In principle, the computation of the thermo- 
dynamic properties from the composition of the 
propellant is a straightforward problem in classical 
thermodynamics. In practice, however, it is the 
developments of the last twenty years which have 
made the solution of the problem possible. The 
development of quantum statistical mechanics and 
the analysis of band spectra has provided the only 
satisfactory method now available for estimating 
the heat capacities of the constituent gases at the 
temperatures as high as those encountered in guns 
and rockets (of the order of 2500 to 4000 K). 
Further, these developments, supplemented by data 
obtained from modern low-temperature calorim- 
etry, have made possible the calculation of equilib- 
rium constants under conditions such that accurate 
direct measurements are experimentally impractical. 

It would be far beyond the scope of this report 
to attempt to outline, in any detail, the process of 
calculating the thermodynamic properties of a 
propellant gas. Such an outline represents a sizable 
manuscript in itself. A schematic block diagram is, 
however, given in Figure 2 and serves to indicate 
the general steps in the process. Rather detailed 
discussion of the methods of building up the requisite 
thermodynamic tables is given in reference 3. The 
reference also provides such tables and carries 
through in detail several examples of the application 
to specific propellants. References 4 and 5 apply 
these methods to detailed calculations for a number 
of other propellant compositions. 

Actually, references 3, 4, and 5 are concerned 
with finding the isochoric flame temperature (of 
interest in gunnery) and the properties of the gas 
under these conditions. However, as indicated in 
Appendix 2 of reference 1, the conversion from the 
isochoric to the isobaric flame temperature is a 
relatively simple matter. 0 

Although the methods of reference 3 are capable 
of considerable accuracy, they are somewhat labori- 
ous, and for this reason simple, more approximate 
methods of estimating the thermodynamic proper- 

0 In references 3, 4, and 5 the conversion is particularly 
simple, since the isobaric flame temperatures are essentially 
the temperatures on the Mollier charts of enthalpy versus 
entropy at which the enthalpies are equal to the “enthalpy 
constants’’ (symbol Hi in reference 3 and 5 and symbol A 
in reference 4) . 


HEAT LOSS, INCOMPLETE REACTION, POWDER LOSS 


75 


ties of the propellant gases have been developed. 
These developments and illustrations of their use 
are described in references 6 and 7. Summaries and 
tables are available in Appendix 8 of reference 1 
and in the Appendix of reference 2. 

It must be remembered that the justification for 
the simple methods of reference 6 is based on agree- 
ment with the more complete methods of reference 
3. In this sense the simple method may be con- 


portions of inorganic constituents) does not lie 
within the methods described but, rather, is due to 
the almost complete lack of adequate basic thermo- 
dynamic and spectral data for these other constitu- 
ents and their reaction products. In this sense, 
thermodynamics is like a large production machine; 
poor raw material leads to a poor finished product, 
and absence of raw material leads to no product 
at all. 



Figure 2. Block diagram of procedure for calculating the thermodynamic properties from the composition 
of the fuel. 


sidered an “interpolation system” for the more 
complete treatment. In the case of the application 
to propellants of composition widely different from 
those previously treated, it would appear to be wise 
to recheck the simple scheme against the complete 
one, modifying the constants of the former as neces- 
sary to bring it into agreement with the latter . 

The success of the thermodynamic calculation 
discussed in this section is essentially limited to 
fuels composed almost entirely of compounds of 
carbons, hydrogen, oxygen, and nitrogen. The 
inadequacy with respect to fuels containing appre- 
ciable quantities of other kinds of constituents (such 
as composite propellants which contain large pro- 


9 6 HEAT LOSS, INCOMPLETE REACTION, 
POWDER LOSS, AND OTHER 
MODIFYING FACTORS 

The preceding sections deal with the theoretically 
ideal performance of a rocket motor. In practice, 
many factors arise which prevent the attainment of 
such ideal performance. Detailed description of 
these factors and their effects would require a 
lengthier dissertation than can be given in this 
report; however, since their recognition and mini- 
mization represent a large part of the science of 
rocketry, a brief summary is presented in Table 1 . 
Most of the information in this table may be in- 


76 


THERMODYNAMIC PROBLEMS 


ferred from the definitions of the quantities involved, 
although in some cases other sources must be 
called upon. 

Some brief comments and qualifying remarks 
may be useful in understanding Table 1. The 
effect of heat loss on the specific impulse and dis- 
charge coefficient arises largely from the lowering 
of the flame temperature, although there is a slight 
effect due to the accompanying small increase in 7. 
The thrust coefficient, being affected essentially 


decrease uniformly. However, when the incom- 
pleteness of reaction is not extreme, it appears likely 
that the deviation of the thrust coefficient from the 
theoretical will be negative and relatively small. 
Increased operating pressures will increase the 
completeness of reaction simply because the gas 
phase reaction proceeds more rapidly at higher 
pressures. Incomplete reaction is generally more 
predominant with “cool” (low flame temperature) 
than “hot” powders. 


Table 1 . Deviations of static* measurements from theoretical values. 

Influence of operating conditions 


Modifying 

factor 

Deviation from ideal 

Specific Discharge Thrustf 

impulse coefficient coefficient 

on deviation 

Operating Initialff 

pressure powder temperature 

Comments 

Heat loss 

- 

+ 

Small 

‘ Generally 
not large 

Generally 
not large 

Uniform from shot to shot 

Incomplete reaction 


+ 

Small, 

probably 

Effect decreases 
with increasing 
pressure 

Generally 

negligible 

Uniform from shot to shot 

Powder loss 


+ 

Negligible 

Increases with 
increasing 
pressure 

Generally marked 
at very high and 
very low 
temperatures 

Erratic from shot to shot 

Poor nozzle 
approach 

— 

— 

— 

Small 

None 

Nozzle approach generally 
badly eroded during shot 

Poor nozzle 
expansion section 


None 


Depends 
on design 

None 

Excessive divergence of 
cone — roughness or poor 
contour leading to non- 
adiabatic flow 

Pressure gauge 
recording high 

None 

— 

- 

None 

None 

Emphasizes the impor- 
tance of frequent gauge 

Thrust gauge 
recording high 

+ 

None 

+ 

None 

None 

recalibrations 

General erratic behavior 

Poorly controlled 
instrumentation 

Variable 

Variable 

Variable 

None 

None 

with little or no correla- 
tion with operating con- 
ditions 


* That is, with rocket motor held in test stand, 
f See qualifying remarks in context. 

tt Operating pressures are generally increased by increasing initial powder temperature. Care must be taken in separating pressure and temperature 
effects. 


solely through the change in 7, is decreased much 
less markedly than the impulse. It may be noted 
that at higher expansion ratios the influence of 
changes in 7 is greater, so that at very high ex- 
pansion ratios (such as might prove useful in very 
high-altitude propulsion) the change in thrust coeffi- 
cient may become somewhat more marked than 
in the case of typical artillery rockets of World 
War II. 

The effect of incomplete reaction is similar to 
heat loss except that increasing incompleteness of 
reaction does not necessarily mean uniformly in- 
creasing 7, so that the thrust coefficient may not 


Powder loss in a given rocket motor increases with 
increasing pressure because of the greater stresses 
thus applied to the charge. Superimposed on this, 
however, there is often a marked increase in loss at 
high powder temperatures (where the pressure is 
generally high) due to the “softening” of the grains, 
and at low powder temperatures (where the pressure 
is generally low) due to increased “brittleness” of 
the grains, which results in tendency to fracture 
under the shock from the igniter. Powder loss does 
not influence the thrust coefficient unless there is 
significant temporary blocking of the nozzle, in 
which case the result is more apt to be a motor 


THE ATTAINABILITY OF HIGH SPECIFIC IMPULSE FUELS 


77 


rupture than a recorded deviation of the thrust 
coefficient. 

Excessive roughness in the nozzle approach may 
decrease the efficiency through skin friction. Sharp 
edges may produce excessive turbulence or a “vena 
contracta” which reduces the effective throat area, 
with the result that best advantage of the expanding 
cone is not obtained. 

Roughness or poor contours in the expanding 
section of the nozzle may lead to the development 
of “shock waves” in the nozzle with lowering of the 
specific impulse. Such difficulties tend to become 
more predominant at higher expansion ratios. 

For the general run of artillery rockets of World 
War II the observed specific impulses ran from 
about 5 to about 10 per cent below the theoretical 
(except in cases of large powder losses) . The devia- 
tions appeared to be largely due to heat loss, al- 
though imperfect nozzle design probably made some 
contribution. Thrust coefficients about 2 to about 5 
per cent low appeared to be the general observation. 
The deviation is again probably attributable to heat 
loss and imperfect nozzle design. 

97 THE ATTAINABILITY OF HIGH 
SPECIFIC IMPULSE FUELS 

It will be noted from the discussion of Section 9.2 
that the principal properties of a fuel which deter- 
mine its specific impulse are the molecular weight 
and temperature of the propellant gas, both of 


which enter the specific impulse as their square 
roots. A typical rocket fuel of World War II might 
have, for example, a flame temperature of 3000 K, 
an average molecular weight of 25, with a specific 
impulse of, say, 210 under ordinary operating 
conditions. 

Consider the possibility of a fuel better by a 
factor of 3 than such a conventional fuel. Suppose 
the improvement were to be obtained by an in- 
crease in temperature. Then a temperature of 
27000 K would be required. Aside from the dif- 
ficulties of finding a chemical reaction to produce 
such a temperature, one can imagine the problem 
of finding materials from which to form rocket 
walls and nozzles, capable of withstanding such 
conditions. 

On the other hand, let the improvement be sought 
in the form of a reduced molecular weight. Hydro- 
gen, with a molecular weight of 2, is the lightest 
gas available to us. On this basis we might expect 
to obtain an improvement by a factor of about 3.5. 
Actually the improvement would be somewhat less 
than this because of the weight of an appropriate 
heater for the hydrogen and the low density of 
liquid hydrogen (see Section 9.1). 

It is apparent, therefore, that a fuel improved by 
a factor of, say, 3 over conventional fuels (which is, 
of course, a sizable improvement) will represent an 
outstanding achievement, whereas improvements 
much greater than this would appear to require 
revolutionary developments in the science of reac- 
tion propulsion. 


Chapter 10 

KINETIC PROBLEMS 

By R. E. Gibson 


101 INTRODUCTION 

T wo very important questions in the design 
and functioning of rockets focus our attention 
on the chemical kinetics of the burning of the pro- 
pellant. Both these questions are connected with 
the equilibrium pressure established in the rocket 
chamber. The first question, one of engineering 
design, arises from the necessity of making rocket 
chambers as light as possible, since they really 
amount to dead load, and any reduction in weight 
of the dead load means a gain in payload or in 
velocity. This puts up to the designer of a rocket the 
question of how to make his rocket motor as light 
as possible and at the same time strong enough to 
withstand any internal pressure likely to be de- 
veloped. Control of the internal pressure is very 
important, therefore, from the viewpoints of effi- 
ciency and safety of design. The second question 
arises from the fact that the thrust of a rocket, and 
hence the acceleration it receives, is given by the 
product of the area of the thrust, the throat coeffi- 
cient, and the internal pressure, F = A t C n P. Since 
A t and C n are substantially constant, we see that the 
internal pressure determines the acceleration of the 
rocket and hence its trajectory and external ballis- 
tics, particularly the value of the gravity drop dur- 
ing acceleration. The second question is, therefore, 
can the internal pressure be controlled within 
tolerance compatible with required ballistic per- 
formance. 

The equilibrium pressure in a rocket chamber is 
determined by a balance between the rate at which 
gas is produced by the propellant and the rate it is 
exhausted through the nozzle. The rate at which 
the propellant produces gas is proportional to the 
area of burning surface and the linear rate at which 
the burning surface progresses. The effect of com- 
position, pressure, temperature, and other environ- 
mental factors on linear rates of burning, therefore, 
takes on a very practical significance. It should be 
mentioned in passing that the term “burning” used 
in this connection should not be confused with burn- 
ing in the sense commonly used, namely, to denote 


interaction of the substance being burned with 
atmospheric oxygen. In the “burning” of solid 
propellants, as it takes place in rockets, atmospheric 
oxygen plays no part, although it has been found 
that the accidental presence of atmospheric oxygen 
may lead to confusing results in experimental 
studies. la The term “burning” when applied to a 
propellant refers to the extremely complex chain of 
reactions which go on when the molecules in the 
system, for example, nitrocellulose-nitroglycerin 
stabilizers, undergo rearrangements to give oxides 
of carbon, water, nitrogen, and small amounts of 
other simple molecular species. 

102 LAW OF BURNING— EFFECT OF 
PRESSURE ON LINEAR RATES 

A grain of propellant burns on all exposed surfaces, 
and the burning surface progresses into the body of 
the grain at a linear rate which is the same at all 
points provided that the powder is homogeneous 
and that external conditions are uniform. This law 
is often called the law of burning in parallel layers. 
The linear rate of burning does, however, depend 
on the pressure of the gas over the propellant, the 
original temperature of the grain, its chemical com- 
position, and, to a lesser extent, on factors which 
will be discussed later. 

Two equations have been used extensively for 
expressing the linear rate of burning of a propellant 
as a function of pressure: 

r = a + bP, (1) 

r = cP n , (2) 

where r is the linear rate of burning, P the pressure 
under which the powder burns, and a, b, c, and n 
are empirical constants. 

Considerable thought has been given to the 
adequacy of one or the other of these equations to 
fit the experimental data. Results for some powders 
are better fitted by (1) than by (2), and for other 
powders the reverse is the cRse. Neither equation 
fits within experimental error over a very large 


7 » 


THE PRESSURE EXPONENT 


79 


range of pressure, but either usually gives an excel- 
lent fit over a range of several thousand pounds per 
square inch. This subject is discussed in several 
reports. 2-6 Several important powders developed 
during World War II exhibit a pressure dependence 
of the linear burning rate that is not well expressed 
by either (1) or (2). 3,9 If, however, we use either of 
these equations to derive a formula for the equilib- 
rium pressure in a rocket motor, we arrive at an 
equation which is not misleading and does bring out 
the role of the various factors involved. 

Equations in terms of both burning rate laws are 
derived in Rocket Fundamentals . 23a Equation (3) 
gives the form corresponding to the burning rate 
law (2) and, being the simpler to follow, is quoted 
here. 

r- -.1/(1 -n) 

p ' < 3 > 

In equation (3), P is the equilibrium pressure, S is 
the area of the burning surface of the propellant, p 
is the density of the solid propellant, p g is the 
density of the propellant gas in the chamber, A t 
is the area of the throat of the rocket, Cd is the 
discharge coefficient of the gas, and c and n are the 
constants in the burning law equation. In Chapter 
12 the effects of other factors influencing the 
steady-state pressure are discussed. 

10 3 KINETIC FACTORS INFLUENCING 
THE EQUILIBRIUM PRESSURE 

Equation (3) shows at once that a stable equilib- 
rium pressure can be generated and maintained 
in a rocket only if n for the propellant is con- 
siderably less than unity. If n is equal to 1, 
1/(1 — n) becomes infinitely large, and any small 
change in one of the factors within the bracket will 
cause an infinitely large change in pressure. A 
rocket could not be designed under such conditions. 
On the other hand, if n is zero, an ideal state is 
reached, because under such conditions the equi- 
librium pressure would vary only linearly with the 
quantities within the bracket. In general, if n is 
between zero and one, a stable equilibrium can be 
reached — the pressure will rise or fall to adjust itself 
to the equilibrium value. The values of n for the 
double-base powders available at the beginning of 
World War II lie between 0.7 and 0.8, close enough 
to unity to raise difficult problems in rocket design. 


If we assume an average value of 0.75 for the n of 
these powders we see that equation (3) becomes 

? (4) 

and that the internal pressure varies as the fourth 
power of the parameters within the bracket. The 
balance is a delicate one, for example, a rise of 10 per 
cent in the area of the burning surface will cause 
the pressure to rise more than 40 per cent. Changes 
in the other variables produce equally drastic effects. 

We have discussed in Chapter 9 the limitations 
placed on A t by port area and loading density con- 
siderations, and we have also shown that Cd de- 
pends on the thermodynamic properties of the pro- 
pellant gas. It is unnecessary to discuss these 
quantities further here except to point out that, 
where n is large, the area of the throat of a rocket 
must be held within very close tolerances and that 
erosion during burning can easily upset the pressure 
balance in the rocket significantly. The quantities 
which concern us most in a consideration of the 
kinetics are n, S, and c. 

104 the pressure exponent 

It will be seen at once that extremely practical 
considerations demand that a good propellant have 
a linear rate of burning which varies as little as 
possible with pressure, i.e., n should be as close to 
zero as possible in equation (2), or b and a should 
be as small as possible in equation (1) . This require- 
ment led at once to two lines of research: (1) an 
empirical study of the effect of composition changes 
on the pressure dependence of the rate of burning 
of a powder and (2) theoretical studies to develop an 
understanding of the burning process with a view to 
isolating the factors that determine n and finding 
out how to control them. The theoretical studies 
progressed to the point where a satisfactory general 
theory of the mechanism of burning double-base 
powder was formulated. The Universities of Minne- 
sota and Wisconsin, Division 8, NDRC, and the 
British investigators made major contributions in 
this field (see bibliography listed in reference 4), 
but it cannot be said that any really useful means of 
reducing the pressure dependence of the rate of 
burning has yet come from these studies. The 
empirical studies which will be outlined later in this 
chapter were more successful, and by 1945 a number 


80 


KINETIC PROBLEMS 


of double-base powders with very low pressure ex- 
ponents over a given range of pressure were dis- 
covered. Of these, powders H-4 (T-2), L 4.8, and 
G117B were the most noteworthy examples. 3 

10 5 THE CONSTANT c 

In addition to depending on the pressure, the 
burning rate of a grain of powder depends on its 
composition, its temperature, and the velocity of 
the gas stream in which it finds itself. The radiation 
falling on the powder also influences the burning 
rate, but, since this effect works by raising the 
powder temperature, it need hardly be considered 
to be an independent one. 

When powders whose compositions differ widely 
are examined, we find that they give values of both 
c and n in equation (2) which are different. If, 
however, one examines a series of powders whose 
compositions do not differ widely — for example, 
manufacturing variations of the same basic formula 
— we find that n may be taken as the same for all 
the powders and the variations in burning rate may 
all be absorbed by variations in the constant c. 
Likewise, change of temperature has little effect on 
n but does change the constant c. 

We may assume, therefore, that manufacturing 
fluctuations in composition and variation in am- 
bient temperature affect the equilibrium pressure 
in a rocket by changing c in equation (2). If n is 
large, then changes in c will produce magnified 
changes in P. This, of course, again emphasizes the 
value of reducing n, but, if such a reduction is not 
possible, every effort should be made to reduce the 
variations of c as a result of composition and tem- 
perature fluctuations. These considerations lead 
again to the need of empirical and theoretical 
knowledge about the effect of composition and tem- 
perature on the burning rates of powders at a given 
pressure. 

10 6 THE AREA OF THE 

BURNING SURFACE 

If a constant chamber pressure throughout the 
entire burning time of the propellant is desired, and 
this is generally required for the most efficient de- 
sign, it will be seen that the area of the burning 
surface of the powder must remain constant within 


very narrow limits. When the propellant obeys 
exactly the law of burning in parallel layers, it is a 
relatively simple matter to calculate the burning 
surface area at any instant if the original geometry 
of the grain is known, and it is possible to arrange 
this geometry in such a way that the burning surface 
does remain constant within the desired limits 
throughout the reaction. A singly perforated cyl- 
inder burning only on the external and internal 
cylindrical surfaces is a simple example of a grain 
whose burning area remains constant, i.e., a neutral- 
burning grain. The cylinder burning on the ends as 
well as the inner and outer surfaces would be a 
•regressively burning grain since the area of the 
burning surface would decrease during the process. 
A number of sufficiently neutral grains were devel- 
oped during World War II. It might be added that 
the increase in port area during burning causes the 
pressure curve to be regressive even for a grain 
having a constant burning surface. In cases when 
this effect is particularly large, it is advantageous 
to have the charge arranged so as to produce an 
increase in surface during burning to give a more 
constant pressure. 

107 RATES OF BURNING OF 

DOUBLE-BASE POWDERS 

The chief experimental work involved in the 
study of the kinetics of the burning of rocket pro- 
pellants consisted of making reliable measurements 
of the linear rates of burning of powders of different 
but known compositions at different pressures and 
temperatures. Other experimental investigations 
concerning the effects of radiation 7,8 and of rate of 
gas flow 4 on the burning rates were also made. At 
the outset of the work the opinion was held by some 
workers with apparent justification that small-scale 
determinations of burning rates were of little value 
in the prediction of the ballistic behavior of a pro- 
pellant in full-scale rockets. However, results of 
subsequent investigations showed that this opinion 
was not well founded and that small-scale ex- 
periments give useful information about propellants 
provided that proper account is taken of all the 
variables involved. An example of the practical 
application of small-scale experiments is to be 
found in the report on the development of a smoke- 
less propellant for the JATO unit. 9 This subject 
is discussed further in Chapter 12. 


RATES OF BURNING OF DOUBLE-BASE POWDERS 


81 


Three distinct methods were used for determining 
the burning rates of powders: (1) closed bomb 
method, (2) vented vessel method, and (3) burning 
strand method. For the same powder, these three 
methods all gave values of the burning rate at a 
given pressure and temperature which were recon- 
cilable, although the task of reconciling them was 
accomplished only after considerable study and con- 
sequent gain in knowledge of the processes involved. 
The methods will now be described. 

10 7 1 Closed Bomb Method 

This method has been extensively used in con- 
nection with gun propellants. A sample of powder 
of known geometry is enclosed in a heavy-walled 


The possibilities of the closed bomb for studying 
rocket propellants were explored at Duke Univer- 
sity and at ABL and several reports are available. 10 
The method is valuable for giving rates of burning 
when the linear mass flow of gas over the propellant 
is essentially zero. In general, however, the closed 
bomb is less useful than the other methods, chiefly 
because its accuracy is best at high pressures and it 
is not well adapted to giving accurate results at low 
pressures — below 2,000 psi, the region of interest in 
rocket work. 

One very interesting phenomenon was observed 
when singly perforated grains were burned in closed 
bombs, namely, that high-frequency vibrations 
were set up during the burning, especially on records 
plotting dP/dT. These vibrations were stopped if a 


RETAINING NUT SHEAR DISK 


NOZZLE 


POWDER GRAIN 

POWDER SUPPORT 


TO PRESSURE GAGES 



LEADS 
IGNITER SQUIBS 


COPPER OR 
MOLYBDENUM 


LARGE MOTOR 

(2 X 12) 


THRUST BEARING 

L 



ADAPTER 


Figure 1 . Small-scale experimental rocket. 


steel vessel, or bomb, capable of withstanding up- 
wards of 100,000 psi. The bomb is provided with a 
water jacket to control its temperature, and with a 
fast-responding pressure gauge by which the pres- 
sure is recorded as a function of time during the 
burning. It is now common practice to use a piezo- 
electric gauge with amplifier and oscilloscope and to 
record pressure and change of pressure with time 
simultaneously. 

In an experiment the bomb is closed tightly to 
prevent gas leakage and the powder ignited. After 
proper corrections for cooling, the maximum pres- 
sure and the change of pressure with time give the 
rate of gas evolution, and this information com- 
bined with a knowledge of the geometry of the grain 
enables one to calculate the linear burning rate at 
any pressure in the region covered. 


steel rod similar to a trap wire in a rocket was 
slipped through the perforation. 10 The phenomenon 
is akin to the “resonance effect” found in rockets 
and mentioned in Chapter 12. 

10 7 2 Vented Vessel Method 

In this method the grain of the propellant is 
burned in an experimental rocket motor fitted with 
a venturi to give the desired equilibrium pressure. 
For any series of experiments several motors and a 
large number of venturis of different sizes are re- 
quired. 4,5,15 In order to cut down effects of high gas 
velocity on the burning rate, the motors should be 
so designed that the free port area is much greater 
than the area of the throat of the nozzle. Special 


82 


KINETIC PROBLEMS 


apparatus was used for extruding the powder into 
suitable grain sizes for this work, and the shapes 
and sizes were carefully controlled by machining 
and measurement. The temperature was controlled 
by conditioning the motor and propellant in a suit- 
able thermostat before firing, and the pressure as a 
function of time was measured by rapidly respond- 
ing Bourdon gauges 12 or strain electronic gauges. 13 
Special precautions for getting rid of the exhaust 




TIME IN SECONDS 










POWDER: REL 1 

« 










S/V = 2.74 











K*, = 618 ‘ 
At/Ad — .090 










































0.10 0.20 0.30 0.40 0.50 0.60 

TIME IN SECONDS 

Figure 2. Typical pressure-time curves obtained 
from burning of powder in vented vessels. 

gases and barricades to confine the results of ex- 
plosions were needed. The pressure-time curves at 
different temperatures and the geometry of the 
propellant grain are the primary data and suffice 
to give the burning rate as a function of pressure 
and temperature. A typical experimental rocket 
motor, a pressure-time curve, and a graph showing 
burning rate as a function of pressure are shown in 
Figures 1, 2, and 3. 

By means of this technique, several hundred 
powders covering a wide range of compositions of 


double-base and composite propellants were ex- 
amined at Indian Head and Allegany Ballistics 
Laboratory. The results are to be found in refer- 
ences 3, 5, and 15. In some cases time permitted 
only the gathering of fragmentary data, and these 
results are to be found in the files of Allegany 
Ballistics Laboratory. 



0.5 1.0 2.0 3.0 5.0 

AVERAGE PRESSURE IN 1000 LB PER SQ INCH 


POWDER COMPOSITION 

A- 68 

NITROCELLULOSE 

5 7.55 

INCLUDING %N 

1 3.2 1 

NITROGLYCERIN 

39.96 

POTASSIUM SULFATE 

1.48 

ETHYL CENTRALITE 

1.0 1 

TOTAL VOLATILES 

1.0 0 

LAMP BLACK 

0.1 0 


HEAT OF EXPLOSION: 1258 CAL PER 6 

BURNING RATE DATA: 


TEMP 

n 

C 

PRESSURE L8 

PER SQ 

IN. 

°c 


I0' 4 

1000 

2000 

3000 

4000 



BURNING RATE IN. PER 

SEC 

50 

0.75 

4.51 

0.79 

1.33 

1.80 

2.23 

25 

0.7 5 

3.86 

0.7 0 

1.18 

1.61 

2.00 

-2 5 

0.8 1 

2.00 

0.5 3 

0.9 3 

L 29 

1 .63 



n 

c' 

U 





0.7 8 

0.6 4 9 

2 29 




Figure 3. The linear rate of burning as a func- 
tion of pressure for a double-base propellant. 


10.7.3 Burning Strand Method 

This method is a new one and was developed at 
the Universities of Wisconsin and Minnesota. lb 
The apparatus consists of a strong steel vessel of 
approximately 300-cu cm capacity capable of with- 


SUMMARY OF EXPERIMENTAL RESULTS 


83 


standing a pressure of 25,000 psi. The lid of the 
bomb supports a framework on which a strand of 
powder about 5 in. long may be supported. In- 
sulated leads through the lid connect with two fine 
wires which pass through the strand, one near the 
top and the other near the bottom. A coating of 
polyvinyl alcohol on the lateral surface of the strand 
prevents it from burning on any but the upper end 



Figure 4. Diagram of apparatus for direct deter- 
mination of the linear burning rate of a strand of 
powder. 


surface. The strand and wire are placed in the 
bomb, the lid fastened tightly, and the whole im- 
mersed in a thermostat . The apparatus is illustrated 
in Figures 4 and 5. Inert gas is pumped into the 
bomb until the desired pressure is reached. When 
temperature equilibrium is attained, the strand is 
ignited at the upper end; as the flame passes each of 
the fine wires an electric circuit is broken, and the 
interval between the breaking of these circuits is 


recorded automatically. At the same time gas is 
exhausted from the bomb at a rate sufficient to keep 
the pressure constant. The length of powder be- 
tween the two timing wires is accurately known, and 
hence the linear burning rate may be accurately 
measured. 

This method of measuring burning rates has great 
advantages. It is direct, it requires very little 
powder for an experiment, and it is rapid. A varia- 
tion of this method uses a bomb provided with a 
window so that the course of burning may be ob- 
served visually or by high-speed photography. By 
this method a large number of experimental pow- 
ders have been investigated, and it should prove 
to be a valuable adjunct to any development or 
manufacturing program. It is most valuable for 
comparative measurements, since the radiation 
effects and the influence of the surrounding atmos- 
phere of inert gas produce results that cannot be 
directly compared with those obtained when the 
powder is surrounded by a fairly thick layer of its 
own combustion products. Reports describing this 
technique and presenting the data on a series of 
double-base powders may be found among the final 
reports from the University of Minnesota lb and 
from the Allegany Ballistics Laboratory. 6 

108 SUMMARY OF 

EXPERIMENTAL RESULTS 

10 8 1 Dependence of Burning Rate 
on Pressure 

For most double-base powders and for some com- 
posite propellants, it was found that the burning 
rate data could be expressed within experimental 
error by either equation (1) or equation (2) — the 
linear or the exponential equations — between 200 
and 500 psi. It seemed that equation (1) gave a 
better fit for some powders while equation (2) gave 
a better fit for others. It is certain, however, that 
both equations must be extended by the addition of 
another pressure dependent term if they are to fit 
data down to atmospheric pressure. ld 

Propellants rich in nonexplosive plasticizers such 
as centralite or triacetin were found to give rate of 
burning-pressure curves that exhibited features 
hitherto unobserved. 3, 5,9 The curves were S-shaped 
and even exhibited maxima. Powders L 4.8 and 
H-5 (see Table 2 of Chapter 13) both showed this 


84 


KINETIC PROBLEMS 


behavior. In this type of propellant there is, there- 
fore, a region of pressure over which the burning 
rate varies very slightly. If equation (2) is fitted to 
the burning rate data in this range, the exponent n 
is found to be very small. It is emphasized that the 
region over which such a simple equation fits the 


In Table 1 the exponents and the corresponding 
pressure ranges are given for these powders and 
compared with those of ordinary double-base rocket 
propellants typified by the T— 1 propellant which 
was available early in World War II. The ex- 
planation of the behavior of powders like L 4.8 or 



Figure 5. Photograph of apparatus for direct determination of linear burning rates. 


complicated burning rate-pressure curve is small for 
these powders, but it is at least 1,000 psi. If, there- 
fore, these propellants are burned in a rocket de- 
signed to develop an equilibrium pressure in the 
proper range, they possess all the advantages of a 
powder with a small n. The equilibrium pressures 
vary only slightly with temperature, surface, throat 
areas, etc. 


H-5 is not complete. It seems true, however, that 
at low pressures the nonexplosive plasticizers do not 
react completely with the explosive ingredient and 
hence the flame temperatures are higher than they 
would be if equilibrium were reached, because more 
carbon dioxide is formed. At higher pressures the 
nonexplosive plasticizers take more part in the reac- 
tion — carbon monoxide is formed in place of carbon 


SUMMARY OF EXPERIMENTAL RESULTS 


85 


Table 1. Burning properties of various double-base 
propellants. 


Temperature coefficient 
Pressure (percentage change in 
Pressure range pressure per degree 
Propellant exponent (n) (psi) centigrade) 


L 4.8 

0.21 

800-1,500 

0.1 

H-5 

0.38 

1,500-3,000 

0.6 

MJA 

0.46 

800-4,000 

0.3 

T-2 

0.69 

1,000-4,000 

0.8 

T-l 

0.73 

1,000-4,000 

1.5 


dioxide, and the flame temperature drops to the 
value expected on the basis of complete combustion. 
Since the rate of burning depends on the flame 
temperature, this explanation does give a picture 
which seems to be qualitatively correct. 

10.8.2 D e p en( j ence 0 f Rate of Burning 
on Temperature 

It was found 15 that the burning rate of a propel- 
lant could be generally expressed as a function of 
temperature by an equation of the form 



or by combinations of (3) and (5) as 



In these equations c' and n and T\ are constants 
whereas T is the initial temperature of the powder. 
It will be seen that the larger is, the less will r 
change with T. A considerable variation in T i was 
found in the variety of powders studied, but no 
convincing generalizations were uncovered. 

In actual rocket practice the variation of equilib- 
rium pressure with temperature is a quantity of 
great significance. This quantity should be as 
small as possible to promote efficiency of design and 
constancy of thrust. It was found that reduction in 
n gave better practical results than increase in T\. 
In the last column of Table 1, the “temperature 
coefficients” for the powders are given in terms of 
percentage change of equilibrium pressure with 
temperature under such conditions that the con- 
stants S, (p — p g ), A t , and Cd in equation (3) were 
held constant. It will be seen that for L 4.8 and 
MJA the temperature coefficients are much im- 
proved over that of the classical powders, as repre- 


sented by T-l. Results for a number of other 
powders are in references 3, 9, and 18. This im- 
provement did more than anything else to make 
smokeless rockets possible for the jet-assisted take- 
off of airplanes. 

10,8,3 Dependence of Burning Rate 
on Chemical Composition 

Examination of the burning rates of a fairly wide 
assortment of double-base powders showed that a 
plot of the burning rates at a given pressure against 
the heat of explosions of the powders (measured on a 
water liquid basis) could be expressed quite well by 
a straight line (see Figure 6) . The higher the heat 
of explosion, the greater is the rate of burning under 



Figure 6. The linear rates of burning of a num- 
ber of double-base powders as a function of their 
heats of explosion. 


comparable conditions. 5,15,23 It is possible to cal- 
culate quite accurately the heat of explosion of a 
powder from a knowledge of its chemical composi- 
tion and a table of constants characteristic of each 
ingredient. 23 We have, therefore, a means of deter- 
mining approximately the burning rate of a powder 
if its composition is known, or, conversely, of speci- 
fying a composition of a powder to fulfill certain 
burning rate requirements. For the most accurate 
work, this relation must be supplemented by ex- 
perimental determinations, but it is a good first 
approximation and proved of great value in design- 
ing propellants for new rockets. It was used with 
effect in designing the H-4 powder charge for the 
115-mm aircraft rocket 14 — probably the most satis- 
factory rocket propellant yet developed — in the 
short space of a few weeks. 


86 


KINETIC PROBLEMS 


As a general rule, it was found that the slower 
the burning rate, i.e. , the lower the heat of explo- 
sion, the smaller was the temperature coefficient for 
a powder. The important composition effect pro- 
duced by the presence of large amounts of coolants 
such as triacetin has been already discussed under 
the dependence of burning rates on pressure. 

10 9 INORGANIC SALTS 

Inorganic salts such as potassium nitrate or 
potassium sulphate are well-known minor constit- 
uents of powders, but their effects on rocket pro- 
pellants were not fully explored until recently. It 
is desirable to discuss these effects in two parts: 
first, the effect of small amounts of salts and, sec- 
ond, the effects of very large percentages of salts. 
When present in small amount (1 to 3 per cent), 
potassium salts modify the burning properties of 
the powder in several desirable ways: (1) they in- 
crease the ease of ignition, (2) they promote regular 
burning at low pressures, (3) they tend to reduce 
flash in the exhaust gases, and (4) they modify the 
course of the pressure-time curve. The flash- 
reducing properties were well demonstrated in the 
use of H-4 (T-2) powder both in the 115-mm air- 
craft rocket 14 and in certain modifications of the 
43^-in. spinner rocket. 11 In general, the elimination 
of flash is brought about by cooling the exhaust 
gases to a sufficiently low temperature before they 
mix with the atmosphere. Two factors assist in this 
cooling process: the use of “cool” powder and the 
use of a large expansion ratio in the rocket nozzle. 
Both these effects, however, are helped by the 
addition of potassium salts to the powder. For 
example, it was found that in a given rocket a 
powder containing potassium nitrate was essen- 
tially flashless, whereas a powder of approximately 
the same heat of explosion but not containing potas- 
sium nitrate gave a brilliant flame in the exhaust. 

The modification of the pressure-time curve by 
the introduction of potassium salts into a given 
powder composition was traced to the effect of 
radiation. 15 The presence of potassium salts in the 
hot gases from a powder increases the emissive 
power of the gases, and hence more radiation falls 
on the burning propellant per unit time, per unit 
thickness of radiating gas. If the opacity of the 
propellant grain is not sufficient to absorb all the 
radiation in a very thin outer layer, radiation will be 


absorbed in the body of the grain, and a rise in 
temperature will result. This will cause an increase 
in the rate of burning and a consequent rise of 
equilibrium pressure in the rocket. Since the 
amount of radiation falling on the propellant grain 
depends on the time, it will be seen that this pro- 
vides a mechanism whereby the burning rate in- 
creases as the propellant is consumed, i.e., the 
powder burns progressively. In the J^-in. singly 
perforated stick granulation, it was found that the 
JPT powder without potassium nitrate gave re- 
gressive pressure-time curves, that is to say, the 
pressure rose to a maximum and then fell off slowly 
as .the propellant was burned. This was due to the 
fact that these grains burned not only on the 
cylindrical surfaces but also on both ends, and 
consequently the area of the burning surface de- 
creased during the reaction. The port area also 
increased during burning. When potassium nitrate 
was added to the composition, progressive pressure- 
time curves were obtained, the pressure rising 
steadily to the end of the burning. 5,15,16 It was 
found that the amount of progressivity in the burn- 
ing of these grains could be controlled not only by 
the addition of potassium salts, but also by the 
addition of varying amounts of carbon black to the 
propellant in order to control its absorption coeffi- 
cients for radiation. This phenomenon is discussed 
in detail in references 5 and 24. It should be em- 
phasized here, however, that in the design and 
manufacture of a first class double-base rocket pro- 
pellant considerable care should be given to speci- 
fying the proper salt content and carbon black 
content to give the desired type of pressure-time 
curve. The hotter the powder, the more attention 
to these details is required. 

When large amounts of inorganic salts were in- 
corporated in double-base powder, together with 
somewhat smaller amounts of carbon or other solid 
reducing agent, entirely new effects were seen, the 
most important being a marked decrease in the 
pressure exponent of the powder. This phenomenon 
is exemplified in solvent-extruded composite propel- 
lants which consisted of a nitroglycerin-nitrocellu- 
lose powder in which was incorporated upwards of 
50 per cent of potassium perchlorate or potassium 
nitrate and several per cent of carbon. A typical 
composition of this propellant is given in Table 6 
of Chapter 13. A large and successful development 
program to make these powders was carried out by 
Division 8, NDRC. Part of this is described in 


THEORETICAL WORK 


87 


joint Division 3 and 8 final reports. 17,18 For a full 
account the Summary Technical Report of Division 
8 should be consulted. 

10 10 BURNING rates and radiation 

Some effects of radiation on the burning of rocket 
propellants have been outlined in the preceding 
paragraphs. Generally speaking, the radiation from 
the hot powder gas influences the burning of the 
powder grain by penetrating below the reacting 
layer and causing a rise of temperature which in 
turn increases the burning rate. The magnitude of 
this rise of temperature increases with the total 
emissive power, itself a function of the temperature, 8 
and the. thickness of the hot gas surrounding the 
grain. It also depends on the absorption coefficient 
of the powder itself in the appropriate regions of the 
spectrum. The effect of radiation in causing the 
progressive burning of rocket powders was inves- 
tigated experimentally and theoretically by con- 
sideration of the above-mentioned factors, and the 
agreement between the results of these two lines of 
attack indicates that the phenomenon is fairly well 
understood. 7 ' 8 An interesting example of the in- 
fluence of radiation was noted in the development 
of grains with long burning times (10 seconds) for 
jet-assisted take-off work. 9 

A very drastic example of the effect of radiation 
on the burning of powders of high calorific value was 
discovered early in World War II. The effect was 
so serious that for a while it was doubted whether 
double-base powder could be used as a reliable 
rocket propellant. In the early days of rocket 
development by Section H, Division 3, NDRC, 
considerable trouble was encountered from the pres- 
ence of cracks, fissures, or other flaws in the powder 
grains. When the grains burned, the flame entered 
these fissures, greatly increasing the burning surface 
area over that predicted and causing a large increase 
in internal pressure with consequent violent dis- 
ruption of the rocket. Research in the manufactur- 
ing process resulted in the overcoming of this diffi- 
culty, but the possibility of fissures being present 
in the propellant grains was so serious that a 
powder was developed which was translucent enough 
to allow positive visual inspection of the grains for 
fissures or flaws. This powder had a composition 
similar to JPT in Table 1 of Chapter 13, had a 
high heat of explosion, and was made in grains that 


were absolutely flawless with a negligible percentage 
of rejects. Nevertheless, rockets continued to blow 
up. A technique for extinguishing powder grains 
before the burning was complete revealed that a 
number of the translucent grains developed numer- 
ous fissures during burning, and hence the area 
of the burning surfaces increased with disastrous 
effects. Figure 7 shows the type of phenomena 
encountered with J^-in. JPT powder. Intensive 
study showed that the effect was due to radiation 



Figure 7. Radiation Assuring of a hot double- 
base powder. Comparison of partially burned and 
unburned grains. 


from the hot powder gases penetrating the powder 
and causing strong local heating when absorbed by a 
trace of dirt or an accidental region of high ab- 
sorbing power. The remedy consisted in introduc- 
ing sufficient coloring matter into the powder to 
absorb the radiation almost completely in the outer 
layers. 1,3,14,16 It was found with all hot powders 
(heats of explosion on a water liquid basis greater 
than 1,000 calories per gram) that stable burning 
in rocket motors is possible only when the powder is 
made sufficiently opaque to radiation. At the 
California Institute of Technology the same diffi- 
culties were encountered with a powder very similar 
in composition to JPN (see Table 1 of Chapter 13) 
and solved in the same way. 

io.li THEORETICAL WORK 

As may be well expected, reaction of double-base 
powder to produce oxides of carbon, hydrogen, 
water, and nitrogen is an extremely complex process. 
It has been attacked theoretically both in this 


88 


KINETIC PROBLEMS 


country and in England, and it is safe to say that a 
theory is now worked out to a point where the 
general processes are qualitatively understood and 
some quantitative predictions can be made. The 
theory is not in a shape where definite simple 
generalizations can be made. 

Very briefly, the theory of burning at moderate 
pressures (of the order of 10,000 to 15,000 psi) 
assumes that the burning reaction takes place in 
three stages: a first-order monomolecular decom- 
position which takes place just below the burning 
surface, a second-order monomolecular reaction 
which takes place in the gas phase close to the 
burning surface, and a branched chain reaction 
which takes place in the gas phase at somewhat 
greater distance from the burning surface. The 
second stage has been referred to as the dark-zone 
reaction and the third stage as the luminous or 
flame reaction. The overall rate-controlling step is 
assumed to be the surface reaction, which in all 
probability is an exothermic decomposition reaction 
involving the formation of nitrogen dioxide. The 
rate of this reaction depends chiefly on the tem- 
perature of the powder very close to the reacting 
surface, and this temperature in turn depends on 
the rate of heat transfer from the hot gas phase 
back to the surface. In the steady state there is a 
steep temperature gradient from the powder surface 
to the flame zone and a steady heat flow across any 
cross section between the powder surface and the 


flame. As the pressure increases, the reaction 
zones become narrower and approach more closely 
to the surface, thus increasing the temperature 
gradient, the rate of heat transfer back to the sur- 
face, and the overall burning rate. The dark-zone 
reaction probably involves the production of alde- 
hydic substances and nitrogen oxide. This reaction 
probably contributes about one-half of the total 
heat and always takes place. The flame reaction 
involves the burning of the aldehydic substances 
and the nitrogen oxide to give carbon monoxide, 
carbon dioxide, water, etc. It always takes place at 
high pressures but may fail to go at low pressures. 
It is interesting to note that failure of the flame 
reaction is apparently closely connected with the 
irregular burning of propellants in rockets at low 
pressures and low temperatures. Attempts to im- 
prove this failing on the part of double-base pro- 
pellants have centered around the use of inorganic 
substances to catalyze the flame reaction. This 
theory, although undoubtedly an oversimplification 
of the actual mechanism, is able to account roughly 
for the observed temperature and pressure de- 
pendence of the burning rate, and reasonable activa- 
tion energies for the various stages can be pos- 
tulated. The fundamental mathematical treatment 
developed by Boys and Corner in England is the 
basis of most of the theoretical work. Further 
details of the theory of the burning of powders may 
be found in references lc, 4, 19, 20, and 21. 


Chapter 11 

STRUCTURAL PROBLEMS 

By R. E. Gibson 


111 INTRODUCTION 

I N ORDER TO FULFILL ITS PURPOSE, a Solid rocket 

propellant must be formed into a given size and 
shape, must be supported adequately in the rocket 
motor, and must possess mechanical properties good 
enough to withstand the stresses imposed upon it 
under firing conditions and during handling and 
storage. During World War II all agencies engaged 
in rocket development expended a considerable 
amount of effort in solving problems connected 
with the design of propellant charges, with the 
methods of making these charges, with the stresses 


8, NDRC, and the reader is referred to the final 
reports of that Division for details. 

Under the general heading of structural problems 
may be included the following subjects: (1) charge 
design; (2) granulation; (3) physical properties of 
propellants. 

112 CHARGE DESIGN 

The propellant charge in a rocket may consist of 
one or more “grains” of powder. The individual 
grains may weigh anything from a fraction of an 


WITHHOLDING NUT 



SQUIB— / IGNITER— 1 ‘—SPACER SLEEVES — J NOZZLE CLOSURE 

Figure 1. Diagrammatic sketch of rocket burning laminated charge of powder disks. 


set up in propellants under working conditions, and 
with the development of propellants whose physical 
properties were adequate to withstand these stresses. 
Both theory and experiment were applied to these 
problems. As a general result, it may be stated 
that the empirical work gave solutions to the more 
immediate practical problems, but a satisfactory 
theory of the solid state of colloidal propellants is 
still to be written. 

This chapter will deal mainly with work done 
by the laboratories associated with Section H , Divi- 
sion 3, NDRC, and will be concerned with double- 
base powders. Much work on the physical proper- 
ties of composite propellants was done by Division 


ounce to several hundred pounds. Each grain, 
however, must be so made that its burning surface 
will remain essentially constant during the whole 
burning period. The reasons for this prime require- 
ment were brought out in Chapter 10 of this report. 
The number, size, and shape of propellant grains 
to be used in any rocket depend on the performance 
required of the rocket and cannot be discussed in 
any condensed form. In Chapter 13 of this report, 
some of the general principles as they relate to 
existent propellants are delineated. An excellent 
discussion of the problem of designing a propellant 
charge for a modern high-velocity rocket is given 
in reference 22, where several generalizations of 


89 


90 


STRUCTURAL PROBLEMS 



1 % BURNEO 



1.5% BURNED 



5% BURNEO 




Figure 2. X-ray spark photograph of propellant burning in superbazooka rocket. Note disk powder grains. 



CHARGE DESIGN 


91 


wide application are made. In this section we shall 
merely refer to a few types of granulations that 
were used and name the reports in which they are 
described. 

Where high accuracy of the rocket is needed, it is 
necessary to guide the projectile until the burning 
is finished. This was found to be a requirement in 
the bazooka rocket, and, in order to keep the 



Figure 3. Effect of temperature on burning dis- 
tance of superbazooka rocket. 


launcher down to a convenient length, a very short 
burning time was essential. Thin-web grains of 
fast-burning powders were used, and, in order to 
have sufficient powder to supply the necessary 
momentum, the charge design called for a number 
of singly perforated grains. 1-4 

In the superbazooka (T-59 rocket) which had a 
higher velocity and greater payload than the 


ordinary bazooka, an entirely new type of charge 
was designed. This charge consisted of a number of 
disks of sheet powder having a web of approximately 
0.06 in. The disks were perforated and held in place 
by a steel rod passing through the perforations and 
anchored at the fore end. (See Figure 1.) Adequate 
port area was provided by progressive reduction 
of the size of the disks, the smallest disk being 
nearest to the nozzle. Figure 2 illustrates this 
charge . The photographs were taken by high-speed 
X-ray photography and actually show various 
stages in the burning of the charge. 5 This charge 
had a very short burning time of the order of 20 
milliseconds — so short indeed that equilibrium pres- 
sure was never reached, and hence the effect of 
temperature on burning time and maximum pres- 
sure was reduced to a very low value. Figure 3 
shows the practical advantages of this charge by 
illustrating the effect of temperature on the burning 
distance. The curve marked HVRG refers to the 
rocket using the charge just described. It will be 
seen that the burning distance of the discharge 
varies very little with temperature when compared 
with other experimental rounds and extremely little 
when compared with the standard M6A3 round. 
For fuller details the reader is referred to the com- 
plete reports. 5,6 

In the M-8 rocket and its modifications (see 
Figure 4), the 115-mm aircraft rocket and its 
modifications, and the 4.5-in. spinner rockets, the 
propellant charges consisted of a number of singly 
perforated cylindrical grains, all the surfaces of 
which were allowed to burn. The grains were sup- 
ported by wires passing through their perforations 
and connected together to form a cage-like trap. By 



GRAINS OF PROPELLANT CHARGE 


Figure 4. Cutaway diagram of M-8 rocket showing a portion of the multigrain propellant charge. 


92 


STRUCTURAL PROBLEMS 


control of powder compositions a wide variation in 
burning time may be realized with this type of 
charge even though the maximum web thickness is 
not greater than half an inch. The ease of manu- 
facture of this type of grain and the fact that, since 
a number of grains are used, statistical methods 
may be applied in specifications and inspection 
constitute the chief advantages of this design. The 
disadvantages arise chiefly from the problem of 


Part II of this volume.) Considerable application 
of this type of charge was found in rockets used for 
towing or pushing demolition charges, a develop- 
ment carried out jointly by Allegany Ballistics 
Laboratory and the Corps of Engineers. 11-15 An 
example of a motor used for towing a detonating 
cable is shown in Figure 5. With a fairly slow pow- 
der, burning times up to 5 seconds may be readily 
obtained with this type of charge. 


Figure 5. Rocket for towing detonating cable. Note form of propellant charge and its support. 



IGNITER ASSEMBLY 


CAGE TRAP 


T-2 PROPELLANT 


NOZZLE 


assembling the propellant charge in the round. The 
details of the design and performance of this type of 
propellant charge are given in reports dealing with 
the weapons in which it was used. 6-10 

The technique of designing propellant charges 
consisting of large single grains of powder, either 
in the form of singly perforated cylinders or columns 
of cruciform cross section, was developed by the 
British and applied in this country very successfully 
by Section L, Division 3, at California Institute of 
Technology whose final reports should be consulted 
for a complete account of the subject. (See also 


113 INHIBITED GRAINS 

The possibility of preventing double-base powder 
grains from burning on certain surfaces by coating 
these surfaces with an adherent layer of noninflam- 
mable plastic has greatly extended the possibilities 
of propellant charge design. Certain compositions 
of ethyl cellulose and of cellulose acetate have been 
found suitable as inhibiting coatings, but a large 
number of other agents have been examined. 16 ' 17 
It has been found quite feasible to restrict burning 
of double-base powder reliably by these means, but 


INHIBITED GRAINS 


93 


the problem of diffusion of nitroglycerin from the 
powder to the coating or of plasticizer from the 
coating to the powder is one which can be solved 
completely only by long-time surveillance tests . At 
present ethyl cellulose compositions give least 
trouble from this source. 

Solid cylindrical grains coated with plastic on the 
cylindrical surface and, therefore, restricted to 
burning on the ends have a neutral-burning geom- 


action is that using a single grain, burning only in 
the perforation and on the end adjacent to the 
nozzle. This charge possesses a great advantage in 
that the powder itself insulates the walls of the 
chamber from the action of the hot gases and per- 
mits the use of light alloys in the fabrication of the 
body of the rocket motor. It offers the most prom- 
ising possibilities for developing high-velocity artil- 
lery rockets — possibilities which were realized in the 



C TRAP 


CHAMBER 


SPACER 


M6N1TER ASSEMBLY 
NO. 24717 


COMPENSATOR 


Figure 6. Photograph of disassembled JATO unit showing propellant charge and support. 


etry and give charges of very long burning time. 
This type of charge was investigated extensively at 
Allegany Ballistics Laboratory (following the lead 
of the British workers) in connection with the devel- 
opment of a JATO unit and of a device for pressur- 
izing a one-shot portable flame thrower. The reader 
is referred to the original reports for details, 19-21 
but Figure 6 gives a general idea of the type of 
restricted burning grains used for the propellant 
charge in the JATO. 

Another important type of rocket propellant 
charge that depends on restriction of burning for its 


scale model of the Vicar, a rocket which carried a 
useful payload at a velocity exceeding 2,600 fps. 22 

In this charge, the outer surface and the fore end 
of the grain are coated with plastic and made to fit 
snugly in the rocket motor. Neutral burning is 
achieved by making the perforation star-shaped in 
cross section so that its perimeter is equal to the 
outside perimeter of the cylinder. Such a charge is 
illustrated in Figure 7 and its development and 
performance are described in references 22 and 23 . 
During 1946 and 1947 Allegany Ballistics Labora- 
tory developed this charge to an advanced stage in 


94 


STRUCTURAL PROBLEMS 


the “ Deacon Rocket,” which carries approximately 
100 lb of propellant, 50 lb of payload, and 50 lb of 
deadload, and attains a velocity exceeding 
4,000 fps. 

SCALE IN INCHES 


0 12 3 4 



Figure 7. Diagram of internal-burning charge 
showing star-shaped perforation. 


114 GRANULATION 

Under this heading come all the problems con- 
cerned with the preparation of reliable grains of 
propellant in the proper shapes, sizes, and types to 
meet the requirements imposed by the applications. 

The general methods used in granulating double- 
base powder are outlined in Chapter 13. Here we 
shall merely refer to one or two outstanding problems 
associated with each method. 

In preparing grains by solvent extrusion, control 
of dimensions and the prevention of warping on 
drying were difficult problems. They were solved 
largely by the efforts of the staff of the Hercules 
Powder Company at Radford Ordnance Works, 
and the reader is referred to reports from this 
organization a for a complete account of the work. 
Fissures appearing in the grains after extrusion also 
gave difficulty and were never entirely overcome — 
the use of carbon dioxide to replace air in the presses 
was suggested by the University of Wisconsin group 
and gave considerable promise. The greater solu- 


bility of carbon dioxide in acetone was the basis of 
this proposal. 24 

A considerable amount of work was expended on 
a study of the “dry” extrusion of solventless 
powder. Since the development and manufacturing 
phases of this subject were thoroughly studied b else- 
where, 25 most attention was given to experimental 
work, die design, studies of flow of plastic through 
dies, effect of composition on the extrudability of 
powder, influence of pressure, temperature, and 
rate of extrusion on the finished product. These 
studies were closely linked with examination of the 
product under ballistic conditions. 26 

In connection with the developments outlined 
in the previous section, extensive studies were made 
of methods of restricting the burning surface of 
propellant grains. This work was based on the 
very important developments made by the British 
workers and indeed was chiefly aimed at adapting 
their methods to powders, plastics, and adhesives 
available in this country. Several satisfactory 
methods of restricting powders were developed, and 
extensive studies were made of the effect of stress set 
up by temperature changes and by shock during 
handling, transportation, and firing conditions. 
This work has continued at Allegany Ballistics 
Laboratory under the Hercules Powder Company 
and has been extended and improved. Details of 
the status at the end of 1945 are to be found in 
reference 16. 

A very significant advance in granulation tech- 
nique was made by Division 8, NDRC, in the 
development of double-base powder which could be 
cast in a fluid state and set up to rigid grains of good 
mechanical properties by storage under proper tem- 
perature conditions. Details of this work can be 
found by reference to the Summary Technical 
Report of Division 8. 

115 PHYSICAL PROPERTIES OF 
ROCKET PROPELLANTS 

Under firing conditions a propellant grain is sup- 
ported by a suitable trap and is acted on by forces 
due to setback, differential pressure, and igniter 
shock. The stresses set up are complicated, and the 
definition of those properties whose quantitative 
expression indicates the ability of a grain to stand 
the stresses is even more complicated. Some work 


To the Ordnance Department. 


b See Chapter 7. 


PHYSICAL PROPERTIES OF ROCKET PROPELLANTS 


95 


was done on the quantitative determination of the 
stresses set up in propellant grains under firing con- 
ditions , 27 but it is emphasized that a great deal of 
experimental work is still needed in this field. In 
parallel, studies were made of the elastic properties 
of double-base propellants, such as Young’s 
modulus and the coefficient of thermal expansion, 
and of the resistance of the powders themselves to 
stresses applied in different ways and at different 
rates with the intention of producing mechanical 
rupture. Several pieces of apparatus were devised 
especially to carry out these experiments, partic- 
ularly to duplicate the rates of application of load 
presumed to exist in actual rockets. Compressive 
strength, impact values, tensile strength, resistance 
to indentation were among the qualities measured. 
The results are not susceptible of generalization in a 


condensed form, and the reader is referred to the 
original reports for the results . 28 

A very important method of determining the 
ability of a powder to withstand the stresses set up 
under firing conditions is a comparison of the pres- 
sure-time curves obtained under static and flight 
conditions. By careful measurements of the velocity 
of a rocket during burning it is possible to calculate 
an acceleration-time curve and with the help of 
auxiliary data to convert this into a pressure-time 
curve. If any discrepancies between the static and 
the flight-pressure- time curves are noted, it is well 
to examine carefully the physical properties of the 
propellant and the nature of its support in the 
motor, since trouble that might develop to serious 
proportions is indicated long before it shows itself 
by disastrous effects . 8 - 9 


Chapter 12 

INTERIOR BALLISTICS PROBLEMS 

By F. T . McClure 


121 SIMPLE BALLISTICS 

is pointed out in chapter 10, the burning law 
l \ for most propellants can be represented, to a 
first approximation, in the form a 

r = cP n , (1) 

where c is a constant characteristic of the propel- 
lant and initial charge temperature, and n is a con- 
stant characteristic of the propellant. 15 Detailed 
discussions of the experimental studies of the burn- 
ing laws for powders are available in references 
1, 2, 3, and 4. 

With this form of the burning law, simple con- 
siderations of the balance of gas production and 
discharge lead to the expression 

i- —.1/(1 - n) 

p ■ < 2 > 

for the equilibrium operating pressure of the rocket 
motor. In this equation, S is the powder surface 
area, c and n are the constants of the burning law, 
A t is the throat area of the nozzle, Cd is the dis- 
charge coefficient of the gas, p is the density of the 
powder, and p g is the density of the gas in the 
motor chamber (usually quite small compared to p) . 

The significance of equation (2) as an illustration 
of the influence of large values of n in magnifying 
the effects on P of small changes in S or c is dis- 
cussed in Chapter 10. 

Although equations (1) and (2) represent satis- 
factory approximations in the case of motors which 
have relatively large cross-sectional areas free of 
powder, they neglect effects which become progres- 
sively more important as cross-sectional area of the 
motor chamber is more completely filled with pow- 
der. Thus, in the design of modern, lightweight, 
high-performance rocket motors, more detailed 
knowledge of the burning law and equilibrium pres- 
sure law is necessary in order to include the im- 

a See Section 5.3.3 for another form of this equation. 

b The use of the symbol n in this chapter as the exponent of 
the burning law must not be confused with its use in Chapter 9 
as the inverse of the molecular weight of the gas. 


portant influence of the so-called “ throat- to-port 
ratio.” The throat- to-port ratio, A t /A p , is the ratio 
of throat area to the cross-sectional area of the free 
space in that part of the motor which contains the 
propellant powder (i.e., the so-called “port area”). 

12 2 INFLUENCE OF THROAT-TO- 
PORT RATIO ON THE 
DISCHARGE COEFFICIENT 

Theoretical consideration of the flow of gas in the 
channels along the sides of the propellant grains 
leads to the conclusion that the discharge coefficient 
will have a small dependence on the throat-to-port 
ratio because of the pressure drop and associated 
gas velocity in the propellant channels. Detailed 
analysis is carried out in Appendix 6 of reference 5, 
leading to the conclusion that the effect may be 
represented with reasonable accuracy by an equa- 
tion of the form 

C D ' = C D [ 1 - 4>(A t /A p y], (3) 

where Cd' is the effective discharge coefficient, Cd 
is the ideal theoretical discharge coefficient, and </> 
is a weak function of they c of the gas, running from 
0.21 at y = 1.2 to about 0.23 at y = 1.4. 

12 3 INFLUENCE OF THE THROAT-TO- 

PORT RATIO ON THE BURNING LAW 

12 3 1 Pressure Drop 

The pressure drop along the propellant channel 
results in a “space average” pressure which is 
slightly less than the head end pressure in the 
motor . The space average burning rate correspond- 
ing to this space average pressure determines the 
rate of gas production in the motor. Correction of 
the burning law for this “pressure drop” effect 6,7 
leads to a burning law of the form 

r = cP 0 n [l - i<KA t /A p )T, (4) 

c Defined in Section 9.2. 


96 


RADIATION 


97 


where r is the space average burning rate, and P 0 is 
the head end pressure, and the other symbols have 
their previous significance. 

12 3 2 Erosive Burning d 

There is still another effect of the flow in the pro- 
pellant channel which is more important than those 
mentioned above. The rate of burning of the pro- 
pellant depends on the velocity as well as the pres- 
sure of the gases flowing over its surface. Higher 
velocities produce higher rates of burning. This is 
made strikingly clear by the observed “tapering 
down to the rear” of partially burned grains, in 
direct contradiction to the effect to be expected if 
pressure alone were the sole determining factor in 
the burning law. A relatively complete and detailed 
experimental study of this problem of erosive burn- 
ing is presented in reference 6, leading to the clear 
conclusion that the basic burning law is much better 
represented by the form 

r — cP n ( 1 + kv) (5) 

than by equation (1). In equation (5) v is the 
velocity of the gas, and k is the so-called “erosion 
constant.” The constant k must be determined 
experimentally for the propellant, and reference 6 
discusses the methods of accomplishing this end. 

Again, in ballistic calculations the space average 
burning rate is the quantity of importance, and 
this can be expressed 6,7 in the form 

r =cP 0 n [l + 0M 2 (A t /A p )], (6) 

where k 2 , the “erosion constant in terms of throat- 
to-port ratio,” can be calculated from the erosion 
constant, k, and the thermodynamic properties of 
the propellant gas. 

12 4 BALLISTIC EQUATION INCLUDING 
THE THROAT-TO-PORT RATIO 

When the throat-to-port effects discussed in the 
previous sections are included in the calculation of 
equilibrium pressures, the equation 

c(p-p,)s[i-fr(^M»mi+o.fife(^AM i 1/(1 7^ 

A,Cd[ 1 -4>{A,/A v y} { 


d See also Section 5.3.2. 


is obtained for the equilibrium pressure at the head 
end of the rocket motor. This equation is discussed 
briefly in reference 6 and in more detail, especially 
with respect to rocket design, in reference 7. The 
experimental design work described in reference 8 
verifies the essential correctness of equation (7) 
and establishes it as a basic equation in the design 
of solid fuel rocket motors. Current reports from the 
Allegany Ballistics Laboratory (now operated by 
the Hercules Powder Company under contract 
with the Navy) lend ample support to this claim. 

Equation (7) clearly demonstrates the influence 
of the throat-to-port ratio on the pressure obtained 
in a rocket. Since the port opens up as the powder 
burns away, this throat-to-port effect decreases 
with time (largely due to the decrease in erosive 
burning) . The effect is thus a regressive one, and to 
obtain constant pressure operation (necessary for 
light-walled motors) it is necessary to design the 
grain with a progressive surface to counterbalance 
the throat-to-port effect. The importance of 
equation (7) in determining the desired surface pro- 
gression is obvious. It is also clear that, in order to 
have adequate information on which to design 
modern solid fuel rockets, it is necessary to know 
the three burning law constants, c, n, and k, as well 
as the thermodynamic properties of the powder gas . 

Although equation (7) takes account of the most 
important factors which influence the equilibrium 
pressure of a rocket motor, there are other factors 
which may produce marked effects under more 
specialized conditions. Two of these factors are 
discussed briefly in Sections 12.5 and 12.6. 

125 RADIATION 

The radiation from the hot powder gases also 
affects the burning rate of the powder, and under 
special conditions may produce extremely large 
effects. The general problem of radiation in the 
rocket chamber and its influence on the burning 
of the powder is discussed in some detail in 
references 9 and 10. 

Although qualitative and sometimes semiquanti- 
tative treatment of special radiation effects (such as 
Assuring, end pressure peaks, and influences of wide 
gas channels and metal walls on burning rate) have 
been possible, complete integration of the radiation 
phenomenon into the ballistics system has not yet 
been, as far as the author knows, successfully accom- 


98 


INTERIOR BALLISTICS PROBLEMS 


plished. Present knowledge, however, indicates 
that such integration is possible if the time and labor 
are made available to do the job. 

12 6 RESONANCE EFFECT 

An unusual effect, not yet completely explained, 
is the so-called “resonance effect.” This phenom- 
enon results in the appearance of greatly increased 
pressures part way along in the burning. These 
peak pressures frequently last for only a short 
period, and then the pressure drops again to the 
normal equilibrium value and the burning process 
continues as though nothing unusual had happened. 
The resonance phenomenon is apparently highly 
specific with respect to powder and motor geometry 
and also operating conditions. Although the phe- 
nomenon can generally be prevented by “breaking 
up” the geometry (such as by putting a metal rod 
down the perforation of a grain), there appears to 
be no way as yet of predicting whether or not a 
given motor design will display the phenomenon. 
This would appear to be a realm in which con- 
siderable advance in knowledge is highly desirable. 

12 7 DRAG OF THE GAS STREAM ON 
THE PROPELLANT 

Two factors contributing to the forces tending to 
cause mechanical failure of the propellant in a 
rocket motor are the acceleration forces, and the 
drag of the flowing gases on the propellant charge. 
The former is easily calculated, but the latter is 
more involved. 

Reference 1 1 provides a simple theory of the drag 
of the gases on the charge and a limited experi- 
mental verification of proposed formulas, which give 
the drag as a function of the cross-sectional area of 
the charge and the throat-to-port ratio. Further 
experimental study of these and other forces on the 
propellant charge are of definite interest to the 
future design of solid fuel rockets. 


12 « HEAT TRANSFER TO THE 
MOTOR WALLS 

The problem of the heat transfer from the hot 
gases to the motor walls is important because of the 
consequences in reducing the strength of the metal. 

Reference 12 contains a theoretical discussion of 
the heat transfer problem, and reference 13 con- 
siders some of the experimental problems associated 
with making significant measurements. (Sec also 
Part II and Chapter 23 of this volume.) 

129 NONSTEADY-STATE ROCKETS 

The ballistic laws discussed in the earlier sections 
of this chapter apply to rockets which operate under 
equilibrium conditions. In very special cases, it 
may be advantageous to operate a rocket motor in 
which the pressure is limited only by the complete 
consumption of the propellant. Design of charges 
for this sort of application is discussed in Section 
11.2 hereof. Such motors, however, appear to have 
a very limited application. Their ballistics and the 
design of such a motor are discussed in detail in 
reference 14. 


12 10 SPECIFICATIONS AND 

TESTING OF PROPELLANTS 

The problem of setting down specifications and 
control testing procedures which will assure that a 
mass-produced propellant will behave as intended 
is a difficult one. It can only be approached from 
the standpoint of a basic knowledge of rocket bal- 
listics. This approach was explored during World 
War II, and considerable success was achieved in 
formulating rational specifications based on scien- 
tific knowledge. Reference 15 discusses this prob- 
lem in considerable detail, using as specific examples 
powders which were standardized during World 
War II. 


Chapter 13 


PROPERTIES OF ROCKET PROPELLANTS AVAILABLE OR DEVELOPED 

DURING WORLD WAR II 

By R. E. Gibson 


131 INTRODUCTION 

I n this chapter we shall give a description of 
the properties of propellants which were either 
available or developed between 1940 and 1945. In 
order to make the chapter as self-contained as 
possible, we shall first gather together the definitions 
of quantities significant in the use of rocket pro- 
pellants, then discuss the various classes of propel- 
lants in terms of these quantities, and, in the case 
of each class, present a table summarizing the com- 
positions of representative members. In discussing 
the properties of these various powders, an attempt 
is made to bring out considerations which are of 
significance in the design of new rockets. The 
chapter ends with a short section suggesting lines 
along which research and development work in the 
field of rocket propellants may proceed in the future. 
The substance of this chapter appears as part of 
one of the Allegany Ballistics Laboratory final re- 
ports. 1 The report was originally written by the 
author as the technical section of a final report from 
the Rocket Propellant Panel to the Joint Committee 
on New Weapons and Equipment. It, therefore, 
includes the work of a large number of agencies 
and is wider in scope than the preceding chapters. 

13 2 SIGNIFICANT CHARACTERISTICS OF 
SOLID ROCKET PROPELLANTS 

13 2 1 Specific Impulse and 

Effective Gas Velocity 

The thrust imparted to a rocket when unit mass 
of powder gas is discharged per second a is a quan- 
tity which is of great interest in rocket design and 
which depends primarily on the thermodynamics 
of the propellant gas, as modified slightly by heat 
losses, and secondarily on the expansion ratio of the 

a The mass of gas discharged per second is equal to the 
mass of powder burned per second when a steady state is 
reached in the rocket. 


rocket nozzle. In ordinary units it may be ex- 
pressed as [lb (force) X seconds] per [lb (mass)] and 
is called the specific impulse. It will be seen that 
the specific impulse multiplied by the mass of pow- 
der burned gives the total impulse, that is to say, 
the momentum, given to the rocket. If the thrust 
imparted to the rocket per unit mass of powder dis- 
charged per second is expressed in common velocity 
units by converting lb (force) to lb (mass), it is 
called the “effective gas velocity’’ of the propellant. 
In ordinary units effective gas velocity = 32.2 X 
specific impulse (32.2 being the acceleration of 
gravity) . 

One of the problems in the development of rocket 
propellant is the search for propellants of greater 
specific impulse, since, it will be noted, the velocity 
increase of a jet-propelled device of given weight 
and carrying a given weight of propellant is almost 
directly proportional to the specific impulse of the 
propellant. 

13 2 2 Burning Time 

The total momentum (mass X velocity) given to 
a rocket device may be conveniently regarded as 
the product of the thrust multiplied by the time the 
thrust is applied (more rigorously, the integral of 
the thrust multiplied by the time). The accelera- 
tions and the mass of gas discharged per second, a 
measure of the blast of the jet, are both propor- 
tional to the thrust, and either may impose upper 
limits on the allowable value of the thrust. The 
time of application of the thrust, i.e., the burning- 
time of the propellant, is, therefore, an important 
engineering variable. The burning time of a rocket 
propellant charge depends on two quantities: (1) the 
linear burning rate of the propellant and (2) the 
distance the burning surface must move as the flame 
consumes the propellant. This latter is commonly 
referred to in terms of the web thickness of the 
powder grains. 

The linear burning rate of a propellant depends 


99 


100 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


primarily on its composition, its temperature, and 
the pressure of the gas over it, and secondarily on 
the radiation falling on it and the rate of gas flow 
over its surface. It has been discussed fully in 
Chapter 10 of this volume. 

13 2 3 Web Thickness 

This introduces the problem of the geometry of 
propellant charges. The “web thickness” is a term 
used to describe the minimum distance through 
solid powder between two exposed or uninhibited 
surfaces. Since burning takes place on all exposed 
surfaces, it will be seen that the burning distance is 
usually one-half the web thickness. Since the rate 
of gas production of the propellant is proportional 
to the burning area (other things being equal) , it is 
important that this area be kept constant within 
narrow limits, which become narrower as the pres- 
sure exponent of the powder rises. All rocket pro- 
pellant charge design is based on the law of burning 
in parallel layers, which enables one to calculate the 
area of the burning surface of a grain at any time 
during its combustion . This law must be obeyed by 
any rocket propellant. Cracks, flaws, or porosity, 
therefore, cannot be tolerated. The problem of 
grain design is soluble in all cases only if powders 
with a wide range of linear burning rates are at 
hand. The problem of propellant charge design is to 
arrange the geometry of the fuel in such a way that the 
burning surface remains essentially constant during 
the complete reaction, and is large enough to produce 
the required thrust, while the minimum distance the 
flame must travel is of the proper length to give the 
desired burning time. This minimum distance is 
closely related to the “web thickness” of the pow- 
der, being equal to it or some submultiple of it. 

1324 Granulation 

The main characteristic which differentiates rock- 
et propellants from gun propellants is the size and 
shape of the individual powder grains. Gun pro- 
pellant grains seldom weigh more than a few 
ounces, whereas rocket propellant grains may weigh 
upwards of 100 lb. Although they may be made in a 
variety of shapes, rocket propellant grains all have 
one characteristic in common: the shape must be 
such as to give approximately neutral burning. One 


of the chief problems in making a rocket propellant 
is that of granulation or forming the propellant into 
the desired size and shape of grain . b 

13.2.5 Overall Specific Impulse 

The specific impulse, as we have seen, is equal to 
the total impulse given to the rocket divided by 
the mass of propellant burned. A quantity of con- 
siderable use in evaluating jet motors is the “overall 
specific impulse” or “impulse-weight ratio” which is 
defined as the total impulse divided by the total 
weight of motor metal parts plus powder. Since the 
metal parts of a rocket motor are generally a dead 
load, the overall specific impulse is a measure of the 
efficiency of the design of the whole unit, and the 
augmentation of this quantity is an important ob- 
jective in present and future rocket work. It will be 
recognized that at least half of the work necessary 
to attain this objective involves the development of 
lighter metal parts and is beyond the scope of this 
report. However, the other half presents the fol- 
lowing problems which must be solved by the 
developers of propellants. 

Burning at Low Pressures 

The weight of the motor increases approximately 
in direct proportion to the internal pressure it must 
stand, whereas the specific impulse increases much 
less rapidly with pressure. There is, therefore, a 
distinct weight advantage to be gained by reducing 
the reaction pressure to a point where engineering 
considerations other than the bursting pressure 
become important factors in motor design. This 
requires a propellant charge whose chemical com- 
position and geometry is such that it burns regularly 
at low pressures and has a low temperature coeffi- 
cient. 0 A low value of the pressure exponent d of the 
propellant is advantageous on both these counts. 

High Specific Impulse 

It is hardly necessary to call attention to the fact 
that a high specific impulse of the propellant is 
needed to get the highest overall specific impulse 
of the rocket motor. 

b This subject is covered in Section 13.3.5 and in Chapter 7. 

c See Section 10.8.2. 

d This exponent is n in equation (2) of Section 10.2. 


CHARACTERISTICS OF SOLID ROCKET PROPELLANTS 


101 


Density of Loading 

It is obvious that the overall specific impulse of a 
given propellant and motor combination will in- 
crease as the amount of propellant per unit volume 
of motor increases, i.e., as the density of loading 
increases, and indeed will reach a maximum when 
the motor chamber is completely filled with powder. 
Limitations on the density of loading are caused 
primarily by the need for a large enough burning 
surface to produce the required thrust and by the 
necessity of providing sufficient port area for the 
gases to travel from one end of the rocket motor to 
another. By all odds the most effective way of 
obtaining a high loading density is to use a cylin- 
drical grain which fills the motor completely and 
burns from one end only. This type of charge 
utilizes all the available space and leaves the whole 
cross section area available for gas flow. Its use is 
limited by the fact that all known propellants have 
too small a linear burning rate to give a large enough 
thrust or short enough burning time in vessels of 
suitable shape. 

Thermal Insulation of Bucket Motors 

The temperatures of all propellant gases are of 
necessity very high, and, when the burning times 
exceed half a second, sufficient heat is transferred 
to the metal parts to reduce their strength consider- 
ably. This raises the dead weight of metal needed 
for safe and reliable performance. Two methods of 
insulating the walls have been tried: the first con- 
sists of applying an insulating coating, usually a 
ceramic, to the interior walls of the chamber and 
has not been very successful; the second consists of 
using the propellant itself as an insulator, and this 
shows great promise. In such a loading arrangement 
the propellant is formed as a perforated thick-walled 
cylinder which fits tightly into the motor. The 
outer cylindrical surface and the fore end are 
treated in such a way as to inhibit burning on these 
surfaces. The combustion takes place in the per- 
foration, and the hot gases impinge on only a small 
portion of the walls near the nozzle. Constancy of 
burning surface is obtained by forming the contour 
of the perforation into a star shape of proper size. 
(See Chapter 11.) This type of rocket offers the best 
promise for high loading density combined with 
light motor weight. The propellant problems pre- 
sented are the granulation of powder into large 


perforated cylinders with thick walls and the re- 
striction of the cylindrical surfaces. 

13 2 6 Rate Control, a New Principle 

Hitherto the rate of evolution of gas by a rocket 
propellant has been governed by the linear burning 
rate under the conditions in the chamber, because, 
by design, the burning surface itself is kept con- 
stant. In 1945 a new principle was explored 
by Division 8, NDRC, whereby the burning surface 
may actually change in area during the combustion 
and thereby the rate of gas evolution is made less 
dependent on the linear burning rate of the mass of 
powder. This has been accomplished by embedding 
in a matrix of the double-base powder strands of 
special powders chosen because of their low tem- 
perature coefficient and low pressure exponent. The 
linear burning rate of these strands determines the 
rate of evolution of gas by the whole mass. e 

13 2,7 Gas Temperature 

For the same expansion ratio and chamber pres- 
sure, the specific impulse of a propellant is roughly 
proportional to the square root of the number of 
moles of gas per pound and to the square root of the 
absolute temperature. High specific impulses are, 
therefore, generally accompanied by high gas tem- 
peratures. These are frequently undesirable, espe- 
cially in long-burning rockets, because of the erosive 
effect on the nozzle. By changing the chemical 
composition, it is theoretically possible to produce a 
propellant with a high specific impulse and a fairly 
low gas temperature. Very little progress has been 
made along these lines up to date, but the problem 
is one of the important ones for the future. 

13 2 8 Chemical Stability 

A rocket propellant must conform to all the spe- 
cifications required of a gun propellant in regard 
to stability under climatic, storage, and extreme 
conditions of use. The specifications, and tests 
to ensure conformity with them, are now well 
established. 

e See Division 8 Summary Technical Report for further 
information on this technique. 


102 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


1329 Sensitivity 

It is desirable to reduce the sensitivity of rocket 
propellants to impact, shock from small-arms bul- 
lets, etc., to a minimum. At present there is prac- 
tically no rocket propellant which is not ignited by 
ride fire. 


13.2.10 Mechanical Properties 

The propellant in a jet-operated motor is subject 
to a variety of stresses during its use . These stresses 
come from differential gas pressure in the motor 
itself and from setback forces arising from accelera- 
tion or from shock during handling. Rates of 
applications of these stresses are, in general, quite 
high, and it is essential that measurements made in 
the laboratory to test the physical properties of 
rocket propellants should be made with comparable 
loading schedules. Although a considerable amount 
of work has been done on the measurement of 
physical properties of rocket propellants, it has not 
yet been established what are the really significant 
measurements to be made. It seems, however, that 
Young’s modulus, the impact resistance, plastic 
flow, and failures in tension and compression all give 
results of practical significance if measured over an 
appropriate range of loading rates. 

133 DOUBLE-BASE POWDERS 
General Description 

The name “double-base powder” was originally 
given to colloidal propellants containing two bases 
or materials capable of self-combustion, namely, 
nitrocellulose and nitroglycerin. It has been ex- 
tended to include all propellants made with nitro- 
cellulose and one or more explosive plasticizers such 
as nitroglycerin, diethylene glycol dinitrate, and 
DINA. f In addition to nitrocellulose and the ex- 
plosive plasticizer, these propellants usually contain 
a stabilizer such as centralite and auxiliary plas- 
ticizers such as centralite, phthalate esters, triac- 
etin, dinitrotoluene, and other compounds of this 
nature which also act as cooling agents. In order to» 
suppress flash and to obtain smoothness of burning 

1 Diethanolnitramine dinitrate. 


at low temperatures, it has been found desirable to 
add 1 or 2 per cent of a potassium salt to double- 
base powders. By adjusting the amounts of nitro- 
cellulose, the physical properties of the colloid may 
be varied over a wide range of toughness and plas- 
ticity, and by varying the amount of explosive 
plasticizer and coolants the flame temperature and 
the burning rate may also be given wide variations. 
Several double-base compositions are shown in 
Tables 1,2, and 3. 

Thermodynamic Properties 

The densities of most double-base powders are 
approximately 1.6 grams per cu cm, that is, about 
0.058 lb per cu in. The isobaric adiabatic flame tem- 
peratures vary from 2400 to 3200 K. The specific 
impulses vary from 235 lb-sec per lb for the powders 
containing approximately 40 per cent nitroglycerin 
and 2 or 3 per cent of cooling agent, to 190 for 
powders containing 20 per cent nitroglycerin and 
approximately 20 per cent of cooling agent. The 
number of moles of gas per gram is about 0.040. 

13 3 3 Burning Properties 

At room temperature (70 F) the linear rates of 
burning of double-base powders vary between 0.4 
and 1.2 ips at 2,000-psi pressure. These figures 
correspond to rates of gas evolution of 0.024 and 
0.071 lb-sec per sq in. of burning surface under 
these conditions. It is a general rule that the hotter 
the powders, i.e., the higher the adiabatic flame 
temperature, the higher the burning rates. It is of 
interest to note that at 2,000-psi chamber pressure 
1 sq in. of burning surface gives a thrust of 4.5 lb 
force with the cooler powder and 16.3 lb force with 
the hotter in motors of appropriate design. 

Until recently the pressure exponents (see Chap- 
ter 9) of all known double-base powders were 
undesirably high, being between 0.7 and 0.8. This 
caused irregularity of burning, forced a reduction in 
loading density, and accentuated the temperature 
coefficient of the chamber pressure and thrust. 
Indeed for a fixed rocket geometry the pressure 
and thrust increased approximately 0.8 per cent 
per degree Fahrenheit. Rather severe practical 
limitations to the use of double-base powder 
rockets arose from this. Indeed for a long time 
double-base powder was ruled out from con- 


DOUBLE-BASE POWDERS 


103 


sideration in connection with JATO units on this 
account. Recently several double-base composi- 
tions have been discovered whose pressure expon- 
ents in the range 800 to 2,000 psi are 0.5 or less. 
When these propellants are used, the pressure and 
thrust of a rocket motor change only 0.2 to 0.3 per 
cent (and even as low as 0.05) per degree Fahrenheit 
over the temperature range —40 to 140 F. One of 
these powders is particularly adaptable to use at 
1,000-psi pressure. Unfortunately all are cool pow- 
ders and have relatively low burning rates. The 
reason for the low pressure exponent of these double- 
base powders is not yet completely understood, but 
1945 experiments on a captured Japanese powder 
give a clue which should certainly be followed. 

13 3 4 Mechanical Properties 

If properly made, double-base powders can be 
obtained as tough, nonporous, homogeneous colloids 
which obey perfectly the law of burning in parallel 
layers. The mechanical strength and elastic prop- 
erties such as Young’s modulus rise rapidly with 
the nitrocellulose content. It should be noted that 
double-base powders colloided with the aid of an 
active solvent are much stronger and tougher than 
those made by rolling and dry extrusion. In general, 
the mechanical and elastic properties of the better 
developed double-base powders are adequate at 70 


F to stand the stress set up during the projection of 
a rocket. At high temperature, i.e., above 100 F, 
experience has shown that these propellants flow too 
easily and have too low a value of Young’s modulus 
to be satisfactory. Furthermore, at low temperature 
their impact strength falls off so rapidly that powder 
breakup from brittle fracture occurs during the 
launching of many rockets. These defects have been 
studied, but, although promising clues have been 
found, a considerable amount of research work is 
necessary to put this aspect of the subject on a 
sound theoretical and practical basis. 

13 3 5 Granulation 

Double-base powder may be made in grains suit- 
able for use in rockets by four different processes, 
each of which has its own advantages and limita- 
tions. These are (1) solvent extrusion, (2) solvent- 
less extrusion, (3) casting, (4) pressure molding. 

Solvent Extrusion 

In this process an active volatile solvent is added 
to the nitrocellulose-nitroglycerin mixture, and the 
whole is stirred in an incorporator. The solvent 
swells the nitrocellulose and permits colloiding, i.e., 
breakdown of the fibrous structure, with a small 
amount of mechanical work. The soft paste or 


Table 1. Nominal compositions of standard double-base powders. 


'''-'--^Powder 

Ingredient "" 

JPT 

JPT 

M13 

T-2 

(H-4) 

H- in. 

Stick 

JPN 

Cordite 

S.C. 

Cordite 

SU/K 

Cordite 

R.S. 

Nitrocellose 

58.80 

57.30 

58.00 

58.25 

51.50 

50.00 

50.00 

57.00 

Per cent nitration 

13.25 

13.25 

13.15 

13.25 

13.25 

12.20 

12.20 

12.20 

Source* 

WP or CL 

WP or CL WP or CL 

WP or CL 

CL 

WP 

WP 

WP 

Nitroglycerin 

40.00 

40.00 

30.00 

41.00 

43.00 

41.00 

41.00 

28.00 

2-4 Dinitrotolucne 



2.5 





11.00 

Ethyl centralitc 

1.00 

1.00 to 3.00 

8.00 


1.00 

9.00 

9.00 

4.00 

Diphenylamine 

0.2 



0.75 





Diethylphthalate 





3.25 




Potassium sulfate 


1.50 

1.5 


1.25f 




Potassium cryolite (added) 






2.25 


Carbon black (added) 


0.05 

0.02 


0.2 




Methyl cellulose (added) 





0.1 




Candelilla wax (added) 





0.075 




Lead stearate (added) 


0.015 







Heat of explosion (water liquid 









basis) cal per gram 

1300 


930 

1316 

1230 

960 

955 

900 


* WP = wood pulp. 

CL = cotton linters. 


t Not included in heat of explosion calculations. 


104 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


dough so formed is extruded through dies of the 
proper size and shape and cut to length. The 
solvent is removed by drying at elevated tem- 
peratures in forced-air-dry houses. Powder granu- 
lated by this method is tougher and harder than 
powder of the same composition granulated by other 
methods. It is, therefore, indicated in cases where 
high accelerations are needed, because its fibrous 
structure helps resist fracture by the setback forces. 
The action of the solvent in reducing the explosive 
power and sensitivity of the paste reduces hazards 
of manufacture. The chief disadvantage of solvent- 
extruded powder is the severe limitation on the web 
thickness imposed by the necessity of removing the 
solvent. It is not feasible to produce this powder 
with web thicker than half an inch because of the 
very long drying time and the production of cracks 
during shrinkage, attendant on the solvent removal. 
Furthermore, exact control of shape and dimen- 
sions in very difficult in solvent extrusion. 

Table 2. Nominal compositions of promising experi- 
mental double-base powders. 


Powder 


Ingredient^~^'~--_ 

H-5 

L 4.8 

G 117 B 

JPH 

Nitrocellulose 

58.00 

58.50 

50.00 

54.50 

Per cent nitration 

13.25 

13.20 

13.25 

12.60 

Source 

WP 

WP 

WP 

CL 

Nitroglycerin 

20.00 

22.50 

30.00 

43.00 

Dinitrotoluene 

2.50 

2.50 

14.50 


Ethyl centralite 

8.00 

8.00 

4.00 

1.00 

Triacetin 

10.00 

8.50 



Potassium sulfate 

1.50 


1.50 

1.50 

Carbon black (added) 



0.02 

0.10 

Lead stearate (added) 

0.40 

0.40 

0.40 


Heat of explosion 





(water liquid cal 
per gram) 

632 

699 

940 

1252 


Solventless Extrusion 8 

In this process the nitrocellulose-nitroglycerin 
mixture is colloided by severe mechanical working 
on heated rolls without the action of a solvent. The 
resulting sheet powder is extruded hot (110 to 170 F) 
through appropriate dies, annealed, and is then 
ready for use. In this process the web thickness is 
limited only by the sizes of press available. At 
present, with the 18-in. press at Inyokern, powder 
grains with cross section areas equivalent to a circle 
9 in. in diameter can be successfully extruded with 


g See Chapter 7. 


webs 3 in. or larger. Exact control of shape and size 
is readily feasible in solventless extrusion, but the 
powder extruded by the solventless process is not as 
tough and strong as solvent powder. The process is 
quite hazardous, heavy and costly machinery and 
barricades being required. It is, however, the most 
important source of rocket propellants now avail- 
able. Examination of German and Japanese pro- 
pellants indicates that they are definitely stronger 
than those produced in this country and presents a 
clue to the improvement of the strength of solvent- 
less double-base powder that should be followed at 
once. 

Table 3. Nominal composition of cast double-base 
propellant. 

Only one powder has been investigated thoroughly enough 
to warrant it being considered for standardization. Its com- 
position is given as follows: 

Matrix 

Casting powder 35 parts by volume 

Casting solvent 12 parts by volume 

Casting powder — granulated in cylinders 0.030 in. diameter, 


0.030 in. long. 

Nitrocellulose (13.15 per cent nitrogen) 74.0 

Nitroglycerin 20.0 

Diethylphthalate 5.0 

Ethyl centralite 1 .0 

Carbon black (added) 0.5 

Casting solvent 

Nitroglycerin 64.0 

Dimethylphthalate 35.0 

Ethyl centralite 1 .0 

Rate control strands 

Nitrocellulose (12.6 per cent nitrogen) 25.0 

Potassium perchlorate (3 microns) 56.0 

Carbon black 9.0 

Ethyl centralite 1 .0 

Plasticizer 9.0 


The plasticizer consists of 74 per cent nitroglycerin , 25 per cent 
dimethylphthalate, and 1 per cent centralite. 0.28 per cent 
magnesium stearate is added to the whole mixture. 


Casting 

The starting material for this process is finely 
granulated, previously colloided powder, such as 
double-base rifle or pistol powder. Cut or ball 
powders are both serviceable. The small particles 
of this nitrocellulose-nitroglycerin powder are mixed 
with a sufficient quantity of an active, nonvolatile, 
casting solvent (e.g., nitroglycerin dissolved in tri- 
acetin) to form a pourable slurry. This is cast into a 
mold which may be a metal container or a plastic 


CAST PERCHLORATE PROPELLANTS 


105 


tube . The latter may serve as a restricting material 
if this is desired. Heating for about a day at 60 C 
causes the mass to set to a tough nonporous grain 
which has entirely satisfactory burning properties. 
Provided that care is taken in the selection of the 
composition, there seem to be no limits to the size 
and shape of grains that may be produced by this 
process, and it is particularly well suited to the 
production of large single grain charges. The cast- 
ing process is also well adapted for applying the 
principle of burning rate control by strands of 
special powder, since the propellant may be cast 
directly around the strands. Cast propellants are 
still in the development stage, but they offer many 
advantages . In addition to those already given , the 
simplicity of the equipment and the cheapness and 
comparative safety of the process may be cited. 

By July 1947, Allegany Ballistics Laboratory, 
operated by the Hercules Powder Company, had 
carried the development of one type of cast double- 
base propellant to the stage where rocket thrust 
units carrying single grain charges weighing more 
than 600 lb were fired successfully in flight under 
conditions of extreme acceleration. 

Pressure Molding 

Molded double-base powder has been produced 
by mixing Western Cartridge “Ball Powder” with 
a few per cent of plasticizer and molding it into a 
large grain by the application of heat and pressure . 
The details of the process have not been published, 
and indeed the whole work is in a fairly elementary 
stage. Its significance has been greatly diminished 
since the development of the casting process. 


Summary 

The granulation processes just described cover 
the field of rocket propellants very adequately. The 
solvent process is useful where thin-web grains 
strong enough for rapidly accelerated rockets are 
desired. The solventless process is well adapted to 
produce large grains whose lengths are large com- 
pared with their diameters — there is, however, an 
upper limit to the diameter. The casting process is 
best adapted to producing grains whose lengths and 
diameters are comparable. It is especially suited to 
the fabrication of large-diameter grains. The larger 
the grain, the more economical is the casting 
process. 

13 4 CAST PERCHLORATE PROPELLANTS 
13,41 General Description 

These propellants are made by mixing intimately 
together finely powdered potassium perchlorate 
and an organic binder in a fluid condition. The mix 
is cast into a mold where it solidifies by thermoplas- 
tic or thermosetting action. More recently am- 
monium perchlorate has been used instead of 
potassium perchlorate to cut down the smoke. As 
organic binders asphalt and oil mixtures, paraplex 
styrene resins, rubber bases, and fusible ethyl- 
cellulose-castor oil mixes have been used with 
success. The great advantage of these propellants 
is the extraordinarily simple process by which they 
are produced and the cheapness of the materials 
involved. It may be noted as a matter of interest 


Table 4. Nominal compositions of some cast perchlorate propellants. 


Powder 

ALT-39 

Galcit 61-C 

MA-70 

MA-142 

Bruceton cast 

Ingredient 

— (Aerojet) 

(Aerojet) 



perchlorate 

Potassium perchlorate 
Ammonium perchlorate 

75.0 

75.5 

75.0 

74.75 

74.5 

Base 

25.0* 

24. 5f 

25. Of 

25. 0§ 

25*6 1| 

Catalyst (chromic oxide) 




0.25 

Carbon black 





0.5 


* Asphalt, Union LT-1 (AMS-C15) 90 per cent, oil (AMS-C3) 10 per cent, 
t Asphalt (AMS-C2) 70 per cent, oil (AMS-C3) 30 per cent. 

t Asphalt, LT-1 (AMS-C15) 42 per cent, paraplex RG-2 38 per cent, dibutyl sebacate 8 per cent, Acrawax C 12 per cent. 
§ Asphalt, LT-1 (AMS-C15) 34 per cent, paraplex RG-2 46 per cent, dibutyl sebacate 8 per cent, Acrawax C 12 per cent. 
|| Permafil 2851 98.2 per cent, tertiary butyl perbenzoate 1.3 per cent, lecithin 0.5 per cent, quinone 0.03 per cent. 


106 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


that these are the only solid propellants that do not 
depend ultimately on nitric acid. It is impossible to 
give in detail the properties of all propellants of this 
type that have been studied, so that attention will 
be focused on three propellants, one of the asphalt 
and potassium perchlorate type, one of the asphalt- 
ammonium perchlorate type, and one of the ethyl- 
cellulose-potassium perchlorate type. Compositions 
are shown in Table 4. 

1 3.4.2 Asphalt-Potassium Perchlorate 
Propellant — Galcit 61-C 

This propellant is made by stirring together finely 
ground potassium perchlorate and a hot asphalt-oil 
mixture, pouring into the motor, which has been 
lined with a layer of asphalt, and allowing to cool. 
Alternatively, it may be cast into a mold, removed, 
and coated with asphalt and tape or some other 
restricting medium. 

Thermodynamic Properties 

The density of this propellant is high, being 1.75 
to 1.82 g per cu cm, i.e., 0.063 to 0.066 lb per cu in. 
The adiabatic flame temperature is calculated to be 
2100 K. There are uncertainties in this calculation, 
and this figure is probably too low. The gases from 
this propellant erode the nozzle severely. The spe- 
cific impulse with chamber pressure of 2,000 psi and 
reasonable expansion ratio is 170 to 180 lb (force) X 
seconds per lb. The number of moles of gas per 
gram is 0.036. The ratio of the specific heats at con- 
stant pressure and constant volume is 1.21. The 
propellants yield a dense white smoke on burning. 
Galcit 61-C, like others of this type, is very stable 
and difficult to ignite. 

Burning Properties 

At 60 F the linear burning rate is 1.5 ± 0.1 ips at 
2,000-psi chamber pressure; this corresponds to a 
gas production of approximately 0.098 lb per sec 
per sq in. of burning surface. The corresponding 
value of the thrust developed by the burning of 
1 sq in. of surface is 17 lb (force). The pressure 
exponent of this powder has not been well investi- 
gated, but it is undesirably high, being about 0.75. 
On the other hand, the temperature coefficient of 
the isobaric burning rate is so low that the variation 
of thrust with temperature in a given rocket is only 
about 0.35 per cent per degree Fahrenheit under 
conditions of use. 


Mechanical Properties 

Over the usable temperature range these powders 
are fairly soft and not brittle enough even at low 
temperatures to be easily fractured by rough 
handling, although cracking from thermal stresses 
at low temperatures has been troublesome. When 
directly supported by the motor walls, the propel- 
lant has adequate strength to withstand the stresses 
encountered in service, but it seems certain that 
over most of the service temperature range the pro- 
pellant is too soft for applications such as radial 
burning where it is supported only at one end. The 
mechanical properties of this propellant determine 
the safe operating temperature limits. At high 
temperatures the material becomes soft enough to 
flow, whereas at low temperatures it hardens to a 
point where shrinkage cracks appear. The improve- 
ment of the physical properties of this propellant 
has been a problem of urgency and led to the 
development of the ethylcellulose and paraplex- 
binders. 


13 4 3 Asphalt-Ammonium Perchlorate 
Propellants 

A number of these propellants have been devel- 
oped with a view to eliminating or cutting down 
the amount of smoke produced by cast potassium 
perchlorate mixtures. These are made in essen- 
tially the same manner as the asphalt-potassium 
perchlorate propellants. They contain in addition 
to ammonium perchlorate and asphalt small 
amounts of other plastics and plasticizers, together 
with chromium trioxide which acts as a catalyst. 

Thermodynamic Properties 

The densities of propellants of this type range 
from 1.52 to 1.56 g per cu cm, and specific impulses 
varying from 150 to 190 are reported. The gas 
contains 0.050 moles per g, and the adiabatic flame 
temperature is given as 1830 K. There is con- 
siderable uncertainty in these figures. 

Burning Properties 

At room temperature the linear burning rates of 
the ammonium perchlorate propellants are much 
lower than those of the potassium perchlorate pro- 


MOLDED COMPOSITE PROPELLANT 


107 


pellants, varying from 0.4 to 0.85 ips at 2,000-psi 
pressure. Some compositions have been made 
vhich burn well at 1,000 psi. With these propel- 
lants, burning at 2,000-psi pressure, thrusts vary- 
ing between 4 and 8.5 lb (force) per sq in. of burning 
surface may be obtained. The pressure exponents 
and the temperature coefficients have not been 
investigated. 

Mechanical Properties 

The mechanical properties are quite similar to 
the asphalt-potassium perchlorate propellants which 
have already been described. 

13-4,4 Ethylcelliilose-Potassiiim 
Perchlorate Propellants 

These propellants are in the advanced experi- 
mental stage but could be developed and applied 
fairly readily. One type is made by mixing a hot 
molten ethylcellulose-castor oil mixture with potas- 
sium perchlorate and aluminum or carbon and 
casting the mix into a suitable mold or vessel. On 
cooling, the mass sets up to a tough solid, which has 
better mechanical properties over a wide tempera- 
ture range than does the asphalt composition . 
Another type is made by mixing the perchlorate 
and aluminum or carbon with the General Electric 
Company’s resin “Permafil,” which can be cast 
at room temperature and hardens without shrink- 
age to a rubbery solid of unlimited temperature 
range. 

These propellants have thermodynamic and burn- 
ing properties very similar to the asphalt-potassium 
perchlorate ones, but with the significant difference 
that the added aluminum or carbon brings the 
pressure exponent down from 0.7 or 0.8 to 0.6, a 
very important reduction. Further investigations 
of the effects of substances like aluminum on the 
pressure exponent are strongly indicated as a means 
of improving this type of propellant. 

13 5 MOLDED COMPOSITE PROPELLANT 

General Description 

These propellants are prepared by milling to- 
gether in edge-runner mills a mixture of ammonium 
picrate, alkali nitrate, and a small portion of a 
resinous binder. The powdery product from the 


mills is forced into grains of the desired size and 
shape by compression molding at about 10,000 psi 
in a large hydraulic press. The grains are cured at a 
predetermined temperature for a fixed time before 
use. Because of the nature of the binder used, 
these grains can be easily restricted by a plastic 
coating, which prevents burning on the inhibited 
surfaces. 

The fabrication of this propellant requires a large 
number of small edge-runner mills, although im- 
proved techniques may probably be developed by 
further investigation. It also requires large presses, 
and a considerable number of these are necessary 
because of the comparative slowness of the molding 
operation. The raw materials for this propellant are 
all currently manufactured in large amounts. 
Molded composite propellants produce considerable 
amounts of white smoke, the quantity being smaller 
with the slower burning compositions which contain 
smaller proportions of alkali nitrate. Although a 
large number of these composite propellants have 
been studied, it has been found possible to cover a 
wide range of properties with four compositions: 
CP 401, CP 404, CP 218B, and CP 492. These 
are arranged in order of increasing burning rate. 
Their nominal compositions are shown in Table 5. 


Table 5. Nominal compositions of certain molded 
composite propellants (OSRD Report No. 5700). 


Powder 





Ingredient 

CP 401 

CP 404 

CP 218B 

CP 492 

Ammonium picrate 

72.0 

54.0 

46.5 

41.0 

Sodium nitrate 
Potassium nitrate 

18.0 

36.0 

46.5 

50.0 

Plastic binder 

10.0* 

10.0* 

74)t 

9.0t 

Zinc stearate (added) 

0.4 

0.4 

0.4 

0.4 


* 5.0 per cent ethylcellulose , 5.0 per cent Aroclor No. 1254. 
f5 per cent buraminc resin, 1.5 per cent Santicizer No. 8, 0.5 per cent 
butanol. 

t 4.5 per cent ethylcellulose, 4.5 per cent Aroclor No. 1254. 


13.5.2 Thermodynamic Properties 

The densities of molded composite propellants 
range from 1.66 to 1.79 g per cu cm or 0.060 to 
0.065 lb per cu in. Reliable estimates of the flame 
temperatures have not been made. The specific 
impulses, measured at chamber pressures in the 
vicinity of 1,000 psi and with the optimum expan- 
sion ratio, lie between 160 and 170 and vary little 
with the composition of the propellant. 


108 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


13 5 3 Burning Properties 

All molded composite propellants burn very well 
at low pressures; indeed 500 to 1 ,000 psi seems to be 
the optimum chamber pressure for these fuels. By 
change of composition a wide range of burning rates 
may be obtained without much change in specific 
impulse. For example, at 1,000-psi pressure and 70 
F the linear burning rates of CP 401 and 492 are 
0.24 and 1.0 ips respectively. The corresponding 
gas production figures are 0.014 and 0.064 lb per sec 
per square inch of burning surface, and this gives 
thrusts per square inch of burning surface of 2.3 and 
10.6 lb (force) . It should be noted that the ratio of 
nitrate to picrate is the principal factor in determin- 
ing the linear burning rate, but the particle size of 
the nitrate is also an important factor in those pow- 
ders which contain potassium nitrate. The pressure 
exponent in the burning rate law for all these com- 
posite propellants is quite low, being on an average 
0.5. This promotes stability of burning at high 
loading densities and gives a very small effect of 
temperature on the pressure and thrust of a given 
motor. Indeed the pressure in a motor charged with 
a molded composite propellant changes only 0.22 
per cent per degree Fahrenheit. 

13.5.4 Mechanical Properties 

When properly made, molded composite propel- 
lants are nonporous solids with a smooth hard 
surface. They obey the law of burning in parallel 
layers. It is essential that the density be controlled 
in manufacture so that it is between 0.950 and 
0.965 times the theoretical fully packed density. If 
the density is below this limit, troubles from 
porosity will arise, whereas, if it exceeds this limit, 
the grains may crack on being removed from the 
mold. 

Compression Strength 

All the composite propellants will withstand 
compressive stresses of 3,000 psi for short times 
even at 60 C and several times this amount at 
room temperature. Since these materials are 
plastics, the value of the compression strength 
depends on the rate of loading, and few laboratory 
measurements under these conditions have been 
made. However, numerous tests of propellant 
grains in rockets subjected to excessive acceleration 


have failed to give any evidence of compression 
failures. 

13 5 6 Impact Resistance 

Molded composite propellants have a very low 
impact resistance; it is about one-tenth that of 
double-base propellants. However, simple shock- 
absorbing mountings made from cork have been 
devised h which enable the propellant grains to 
stand up against any rough usage tests, such as 
dropping on concrete, which do not damage seri- 
ously the metal parts of the rocket motor. 

13 5 7 Thermal Shock 

The resistance to thermal shock leaves something 
to be desired. It depends on the size of grains and 
the severity of the temperature change, and is in 
the state where it is quite advisable to examine the 
effects of thermal shock on any new rocket loaded 
with a molded composite propellant. The chemical, 
thermal, and explosive stability of all composite 
propellants of this type is extremely high, and the 
impact sensitivity is low. 

13 58 Granulations 

The pressure-molding process works best when 
the diameter of the grains is approximately equal 
to the length. Grains whose lengths are much 
greater than their diameters must be produced by 
the cementing together of one or more smaller 
grains. Since adequate cements are available this 
condition introduces no great difficulties and makes 
possible the production of a wide variety of shapes 
and sizes. Up to 1946, grains varying from 1 
to 12 in. in diameter and from 1 to 51 in. in length 
had been successfully made. The only limit to the 
diameter is the size of the press available. 

13 6 SOLVENT-EXTRUDED COMPOSITE 
PROPELLANTS 

These propellants consist of a filler composed of 
carbon black and either potassium perchlorate or 
potassium nitrate dispersed in a binder of double- 


h See Division 8 Summary Technical Report. % 


RECOMMENDATIONS FOR FUTURE WORK 


109 


base powder. The proportions are usually 65 per 
cent filler, 35 per cent binder, although for some 
purposes where reduction of smoke is important the 
fraction of filler has been reduced to 9 per cent. 

In addition to apparatus for grinding the per- 
chlorate or nitrate the equipment needed for making 
these powders is the same as for making solvent- 
extruded double-base powders, and the same limita- 
tions of web thickness apply. 

The great advantage of solvent-extruded com- 
posite propellants lies in the small value of their 
pressure exponent which is approximately 0.45 and 
which permits high loading density and cuts down 
the temperature coefficient of pressure and thrust. 
The specific impulses are about the same as those 
of double-base powders, and the compositions may 
be adjusted to cover a wide range of burning rates. 
Indeed, extruded composite propellants with burn- 
ing rates faster than are feasible with double-base 
powders are readily obtainable. The granulation 
limitations described under solvent-extruded double- 
base powder apply to solvent-extruded composite 
propellants, and their main use is limited to rela- 
tively fast-burning rockets or to the rate control 
strands which are used in conjunction with cast 
double-base powder grains. For this latter purpose, 
strands of composite propellant are ideal because of 
the small effects of pressure and temperature on 
their burning rates, and because of the wide range of 
burning rates that may be realized within the com- 
position scope of solvent-extruded composite 
strands. 

137 PLASTIC PROPELLANTS 

This type of rocket propellant has been developed 
by the British and is quite similar in composition 
and ballistic properties to the American molded 
composite propellants. The main difference lies in 
the binder, which is more fluid and present in larger 
amounts, so that the plastic propellant does not set 
up to a hard mass but retains a putty like consistency. 
It is molded directly into the rocket motors under 
fairly low pressure in the form of central-burning 
charges inhibited on the outer surface by the motor 
walls, a reliable bond between the plastic propellant 
and the steel wall having been developed. The 
puttylike consistency of the propellant allows it to 
expand or contract with the motor wall without the 
setting up of stresses large enough to cause rupture 
or cracking. 


Table 6. Nominal compositions of some solvent- 
extruded composite propellants. 


Powder 

EJA 

EJB 

MJA 

T-4 

Ingredient 





Nitrocellulose 

21.00 

42.00 

26.00 

54.60 

Per cent nitration 

12.66 

13.10 

13.10 

13.15 

Nitroglycerin 

13.00 

26.50 

21.50 

35.50 

Ethyl centralite 
Potassium 

1.00 


2.50 

0.9 

perchlorate 
Potassium nitrate 

55.50 

25.50 

43.00 

7.80 

Carbon black 

9.00 

4.20 

7.00 

1.20 


The thermodynamics and burning properties of 
this propellant are similar to those of the slower 
burning molded composite propellants. 


138 RECOMMENDATIONS FOR 
FUTURE WORK 1 

The foregoing review of the status of solid rocket 
propellants suggests strongly certain lines along 
which future work should proceed and makes 
possible several general recommendations which will 
be advanced in the following. It should be noted 
that these recommendations are of a general nature 
and are independent of any programs that might 
already be planned for the development of specific 
devices. Future work to be undertaken falls natur- 
ally into two classes: 

1 . Development work, which includes the improve- 
ment of existing propellants and especially the 
development to an entirely satisfactory state of a 
few solid propellants which cover the range of fore- 
seeable requirements . The scale of this type of work 
is on a pilot plant or higher level, and its main 
object is to render available to the United States 
reasonably satisfactory propellants which may be 
prepared in quantities at short notice. 

2. Research work. This includes work on a labo- 
ratory level that is designed (a) to make radical 
improvements in existing types of propellants, and 
(b) to broaden the whole basis underlying the art of 
propellant manufacture. 

The discussion in this section will be classified 
according to the types of propellants considered. 

* These recommendations were made early in 1946. Many 
of them were put into effect during preparation of this volume. 


110 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


Solvent-Extruded 
Double-Base Powders 

It is recommended that no further development 
work be conducted on powders of this type since 
three satisfactory powders are now available; namely, 
T-l powder, the T-4 (BBP), and the T-2 powder. 
Specifications for these have been written, and the 
only problems that require consideration are those 
dealing with improvement of the control in large- 
scale production. 


Solventless Double-Base 
Powders 

Compositions which cover the whole range of 
calorific values or burning rates have been inves- 
tigated for this type of powder. It is recommended 
that work leading to the development of three 
compositions which are satisfactory from the ballistic 
and manufacturing points of view be undertaken at 
once. It is further recommended that these powders 
be based on JPN, the high-calorific powder, G 117B, 
medium-calorific powder, and L 4.8 (note: the com- 
positions of these powders as known at present are 
given in Tables 1 and 2). These powders cover the 
range of burning rates obtainable with double-base 
propellants, and the problems connected with their 
manufacture are known to be soluble. Considerable 
improvement in the manufacturability should be 
sought, but no sacrifice in ballistic qualities such as 
smoothness of burning and small temperature coeffi- 
cient should be made. It is emphasized that these 
problems are fairly short range in nature, but they 
should not be regarded as solved until the results 
have been tested on a large scale, since quantity 
production is an important object. 

It is also recommended that immediate steps be 
taken to use existing lines of evidence to improve 
the mechanical properties of these powders, par- 
ticularly the resistance to load at high temperatures 
and the “brittleness” at low temperatures. 

It is also recommended that studies be made of 
the effect of newly developed stabilizers in extending 
the safe life and cutting down the gas production in 
double-base powders. The gas production in these 
powders is now thought to be the major cause of 
cracking during high-temperature storage, an effect 


which at present imposes serious limitations in the 
use of double-base powder in large web grains. 

Problems concerned with the extrusion of solvent- 
less powders in very large grains, for example, 6 to 
10 in. in diameter, should receive high priority. 
This is particularly true for grains restricted on the 
outer surface and having a star-shaped perforation, 
since this type of grain gives the highest promise of 
realizing the largest overall specific impulse in 
rocket motors. The need for motors with high 
impulse and low weight for the launching of guided 
missiles and similar devices becomes more urgent 
every day. 

13.8.3 £ agt D ol ,M e -B age Propellants 

This development, particularly with the use of 
rate control strands, is regarded as one of the most 
promising in the whole field, and it should be pur- 
sued vigorously, particularly in view of the increas- 
ing demand for jet-operated thrust units of larger 
and larger size. It is suggested that attention be 
given to the development of approximately three 
compositions or combinations of compositions in 
burning rate and rate control strands covering the 
same range as that indicated in the solventless- 
extruded powder field. Attention should also be 
given to the development of large radial-burning 
grains with a low temperature coefficient, produced 
either by adjustments of the composition of the 
powder or by the use of rate control devices. The 
preparation of cast double-base grains in very large 
sizes or with star-shaped perforations should also 
receive early attention. In this connection the 
development of a smokeless composition with the 
mechanical properties and adhesive qualities of the 
British plastic propellant would fill a pressing need. 

13.8.4 Pressure Molding of 
Double-Base Powder 

It is recommended that very low priority be given 
to this type of development in the future, since the 
casting process is simpler and leads to the same 
results. Furthermore, the experience of the years 
1942-45 does not justify optimism concerning 
further work on pressure molding of double-base 
powder. 


RECOMMENDATIONS FOR RESEARCH WORK 


111 


13.8.5 C as t Perchlorate Propellants 

The cheapness and availability of these propellants 
and the possibility of obtaining high loading den- 
sities suggest very strongly that development work 
to improve them should be pursued vigorously. 
Careful attention should be paid to the chemical 
engineering problems arising in the manufacture so 
that a more uniform product may be obtained. 
Higher mechanical strength and a wider usable 
temperature range are important objectives that 
should be sought. Improvement of the exponent in 
the burning law is also a very necessary develop- 
ment; clues to this already exist in the action of 
aluminum in some of these propellants. It is also 
recommended that attempts be made to increase the 
burning rate of the smokeless ammonium perchlorate 
propellants or to develop other cast compositions 
with high burning rate and low smoke. 


13.8.6 Molded Composite Propellants 

These are in a fairly well-developed state, the 
only problems really requiring attention being 
improvement of the manufacturing process, better 
control of the uniformity of the product, and 
removal of any instability at high chamber pres- 
sure. It is felt that the field covered by molded 
composite propellants can, in general, be covered 
by others that are more promising in their properties 
or easier to make. Hence it is not recommended 
that an extensive development program be con- 
ducted on this work. Since molded composite pro- 
pellants are the only ones now available for large 
thrust units designed to give very high thrusts, it is 
strongly recommended that facilities for making 
this propellant be kept in working order until a 
completely satisfactory replacement has been 
developed. 


13 8,7 Solvent-Extruded Composite 
Propellants 

At present these propellants with their low ex- 
ponent in the burning law and high rate of burning 
are the best known for rate control applications. 


The development problems of the solvent-extruded 
composite propellants consist mainly in the se- 
curing of positive manufacturing control and 
should be pushed to a point where satisfactory 
specifications for manufacture and quality may be 
written. 


13 8 8 Plastic Propellants 

These are receiving attention in Great Britain, 
and, since they have the ballistic properties of the 
molded composite propellants and there is a pos- 
sibility that the same mechanical properties may be 
developed in cast double-base propellants, it is 
recommended that little work along this line be 
done until the possibilities of plastic cast double- 
base propellants are more thoroughly explored. 


139 RECOMMENDATIONS FOR 
RESEARCH WORK 

The research program recommended recognizes 
two main objectives: 

1. Radical modifications and improvements of 
existing propellants; for example, replacement of 
nitroglycerin by a new explosive plasticizer. 

2. The production of entirely new types of pro- 
pellants with different bases; for example, use of 
high polymers other than nitrocellulose. 

These new propellants will, of course, recommend 
themselves because of outstandingly good physical 
or burning characteristics, or great ease or flexibility 
of manufacture. Such a research program must be 
guided by an understanding of the fundamental 
characteristics involved; namely, the mechanism of 
burning, the control of the physical and mechanical 
properties, and the knowledge of the desirable prop- 
erties of new ingredients and methods of making 
them. 

The research program suggested here, therefore, 
falls into three classes, which not only subdivide 
the research problems into natural groups, but also 
indicate an organization for carrying them out. In 
the following pages this program is set out, first in 
summary and secondly in more detail, so that the 
reader may comprehend its scope more readily. 


112 


PROPELLANTS AVAILABLE OR DEVELOPED DURING WORLD WAR II 


13 10 SUMMARY OF 

RESEARCH PROBLEMS IN THE FIELD 
OF ROCKET PROPELLANTS 

13.10.1 Q agg Theory of the Burning 
of Rocket Propellants (Kinetics 

of Powder Reactions) 

The objective of this class of problem is an under- 
standing of the relations between those quantities 
of significance in the burning of rocket propellants 
and those quantities which may be controlled in 
their manufacture. A satisfactory theory should 
enable one to predict the burning properties of a 
propellant from its composition and to make pow- 
ders with burning properties specially adapted to 
certain purposes. A sound theoretical basis is of 
utmost importance in the guiding of work in the 
whole program. 

13.10.2 Q agg g Physical Theory of the 
State of Colloidal or Other Solid 

Propellants (Statistical Mechanics 
of Solid Propellants) 

This general class of problems is concerned with 
the relation between the molecular properties of the 
ingredients (chemical nature, degree of polymeriza- 
tion) and the physical and mechanical properties 
of the resultant mass. Practical questions, such as 
extrudability, strength, “degree of colloiding,” con- 
trol of soundness and to some extent of burning 
properties, fall into this class. It is highly probable 
that organic high polymers of one sort or another 
will continue to form the basis of solid propellants 
for some time to come. The search for a complete 
understanding of the relationship between the 
characteristics of the molecules and the properties 
of the solid or liquid state in highly polymerized 
systems is one of the most vital physicochemical 
problems of the day, and one which links up the 
study of propellant explosives with that of other 
plastics. 

13.10.3 Q agg Fundamental Develop- 
ments of New Propellants (Chemistry 

of Propellants) 

The work covered by this class depends for its 
success on close coordination with the work listed 
under classes A and B, because these classes cover 
fields nearer to the ultimate application. Class C, 


however, is sufficiently varied and specialized to 
merit separate consideration. In this class is con- 
sidered not only the chemistry of old or entirely 
new powder ingredients, but also the search for new 
ways of restricting the burning of solid propellants, 
the development of semisolid propellants, the de- 
velopment of fuels with low flame temperatures but 
high specific impulses, and the development of pow- 
ders with improved ignitibility. 

13.10.4 More Detailed Outline of 
Program 

Class A. Kinetics of Propellants 

1. Theory of burning of solid propellants. 

a. Study of reaction in solid. 

b. Study of reaction in gas phase. 

c. Influence of environment, pressure, tem- 

perature, radiation on kinetics (rate 
and exponents, etc.). 

2. Thermodynamic studies of powder and powder 
gases . 

a. Specific heat measurements. 

b. Heats of reaction. 

c. Thermal conductivity of propellants. 

d. Temperature measurements near reac- 

tion zone. 

3. Isolation and identification of intermediate de- 
composition products. 

4. Laboratory studies of kinetics of intermediate 
reactions, i.e., reactions in which the known inter- 
mediate products take part. 

5 . Application of theory to specify desirable pow- 
der ingredients and tests of predictions. 

6. Effect of powder composition and burning 
properties. Under this subhead we include all 
studies of types A1 to A4 as applied to nitrocellulose 
powders, composite propellants, and powders with 
entirely new bases and plasticizers. 

7. Development of new techniques for studying 
the kinetics of reaction of gaseous, liquid, and 
solid propellants, for example, application of mass 
spectrograph, radioactive tracers, and high-speed 
photography. 

Class B. Physical State of Propellants 

1. Systematic studies of physical properties of 
existing and new propellants over range of pressure, 
temperature, and rate of application of stress. 

2. Systematic study of plastic properties over 
ranges of pressure and temperature. 


SUMMARY OF RESEARCH PROBLEMS 


113 


3. Studies of molecular characteristics and fun- 
damental chemistry of cellulose, nitrocellulose, and 
other high polymers. Structure of solid propellants. 
Relation of molecular characteristics, e.g., molec- 
ular weights and polar groups, to the properties of 
the solid, e.g., degree of colloiding of powder. 

4. Effect of mechanical working and other ex- 
ternal effects on molecular characteristics and 
structure of propellants. 

5. Study of molecular interaction of plasticizers 
with propellant bases. Influence of bonding on 
physical state of the solid propellants. 

6. Fundamental studies of the adhesion of solid 
propellants to metals, plasticizers, etc. 

7 . Development of new apparatus and techniques 
for studying the molecular properties and the mac- 
roscopic structure of solid fuels. 

Class C. Chemistry of Propellants 

1. Synthetic organic chemistry as applied to 
explosive bases. 

2. Synthetic organic chemistry as applied to 
explosive and nonexplosive plasticizers. 

3. Use of new bases and plasticizers to obtain 
propellants with higher specific impulses but low 
flame temperatures. 

4. Studies of new stabilizers and their action. 
Reduction of gas formation. Improvement of high- 
temperature properties. 


5. Use of ingredients to promote ignitibility of 
powders. 

6. Exploration of new manufacturing methods, 
including entirely new colloiding processes. 

7. Development of semisolid propellants — thick- 
ened monofluids. 

8. Development of restrictive coatings and meth- 
ods of application. 

9. Investigation of thermodynamic and thermal 
properties of powders and powder constituents. 
This is particularly important in the case of new 
constituents. 

10. Application of new methods to chemical and 
physical analyses of powders. 

13105 General 

The program just outlined is given in fairly 
general terms, but it covers the avenues of inves- 
tigation that now' seem worth following and pro- 
vide fairly well-defined objectives. The details 
should, of course, be filled in more completely by 
those who are to undertake the work. It is suggested 
very strongly that one of the first steps to be taken 
by those undertaking the job should be the prepara- 
tion of a monograph giving the present status of 
solid rocket propellants. In this way the outstand- 
ing problems will be brought into sharp relief. 


j In early 1946. 
























































































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: S\ f 




PART IV 


ROCKET WEAPONS AS DEVELOPED AND USED 
IN WORLD WAR II 

By C. TV. Snyder a 


A significant DEVELOPMENT of World War II 
was the resurgence of the artillery rocket as a 
major weapon. This is strikingly illustrated by the 
fact that in 1941 the U. S. Navy had no rocket 
weapons and evinced little interest in them, whereas 
in 1945 the Navy was spending on them $100,000,000 
a month — more than on all its other types of ammu- 
nition combined. 

All the Navy’s rocket weapons, as well as a con- 
siderable portion of those used by the Army, were 
developed by OSRD’s rocket project at the California 
Institute of Technology, Contract OEMsr-418. The 
CIT work began in September 1941, expanded 
rapidly, and continued intensively into late 1945. 
Many reports were issued during this period. Two 
monographs and seven final report volumes on 
rockets, prepared under the contract, recapitulate 
the principal results and conclusions of four years 
of high-pressure activity. b 

The following chapters attempt to provide an 
introduction to these volumes, and to summarize 
them in part, primarily for the benefit of those 
who may be concerned with rocket research in the 
future. 

Since one of the major aims is to explain why CIT 
rockets evolved as they did, certain basic factors 
are given here in the beginning. They are 

1. Propellant. The only rocket propellant which 
could be made available in sufficient quantities to 
meet the requirements of an artillery weapon was 
ballistite. It was far from ideal for the purpose. 

2. Simplicity. The keynote of all designs was 
simplicity. From the beginning the group set for 
itself the task of developing to the utmost the 
simplest kinds of rockets, which could be made in 
enormous quantities cheaply and quickly, in the 
belief that this course was more likely to contribute 

a Assistant Supervisor, Section I (Rocket Design and Devel- 
opment) of Contract OEMsr-418 at the California Institute of 
Technology. 

b These items head the list of OEMsr-418 reports in the 
general bibliography in the appendix. 


to winning the war than more ambitious and com- 
plicated long-term developments. A comparison of 
the little 4.5-in. 29-lb barrage rocket and the fear- 
some V— 2 as to their relative effects on the outcome 
of World War II will show that this conviction has 
been vindicated. 

3. Safety. It was always insisted that the designs 
be thoroughly safe and dependable. This was done 
not only with a view to preventing casualties among 
our own men, but also because of a realization that 
rockets were new to the Services and a poor showing 
at the beginning might prejudice their users against 
them and seriously retard their growth into a sig- 
nificant factor in the victory. 

Experimenters who come afterward, who have 
access to many kinds of propellant with diverse 
properties, who have time to tackle problems of 
greater difficulty and solve them with greater ele- 
gance, and who, having customers eager for their 
products, may be able to design to smaller safety 
factors in the interest of obtaining the last ounce of 
performance, will certainly do things differently. 
This fact should be kept in mind in reading the 
following pages. 

The author joined the rocket group in June 1942, 
just as the first American rocket was starting into 
combat. Asa member, and later an assistant super- 
visor, of the projectile group, he had first-hand 
experience with most of the rockets discussed in 
these chapters and hence can reasonably hope that 
most of what he has written is true. Nevertheless, 
because of the pressure under which these chapters 
had to be written and the unavailability of people 
and information after the development activities 
ceased, this summary is much more of a one-man 
job than is desirable for a work of its kind. It is 
therefore hoped, but not expected, that the number 
of errors may be few and that the subjects which 
the author was not directly concerned with during 
World War II may have been given their proper 
space and emphasis. 


115 




































































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Chapter 14 

MILITARY NEEDS WHICH ROCKETS CAN MEET 


By C.W. Snyder 


141 GENERAL CHARACTERISTICS AND 
USES 

D uring world war ii, short-burning, solid fuel 
rockets were developed to meet many tactical 
needs. It appears that the field for military 
application of such rockets has been fairly well 
explored. Most of the applications for which rockets 
have been found advantageous have been rather 
specialized; solid fuel rockets have supplemented 
shells and bombs rather than displaced them. 
Before a rocket is chosen or designed for any appli- 
cation, therefore, it should be established that the 
rocket promises definite advantages over other 
types of projectiles. For many common tactical 
situations it does not. For some others, rockets 
may be the only possible answer or the most effec- 
tive one. 

All actual or proposed uses of rocket projectiles 
which have come to the attention of the writer 
involve the familiar functions of shell and bombs, 
namely, the delivery of materials to the enemy, 
sometimes at velocities adequate for penetration of 
his defenses. In addition to solid shot, the materials 
carried have included high explosives, chemical 
agents (gas, smoke, incendiary mixtures, etc.), 
illuminating flares, and certain inert fillers like anti- 
radar “window” and propaganda leaflets. 

The principal characteristics of rockets which 
affect their employment are 

1. Their lack of recoil. This is unquestionably 
their most important advantage and is a factor in 
nearly all tactical uses. 

2. Simplicity, light weight, and associated mo- 
bility of rocket launchers as compared to guns. 

3. Low setback forces resulting from the usually 
prolonged period of propulsion. 

4. Long, stable underwater and underground 
trajectories in the case of most fin-stabilized rockets. 

5. Superior accuracy and penetrating power of 
rockets as compared to bombs. 

Among the characteristics of rockets which have 
limited their use are blast, smoke (in some cases), 
and the effects of temperature on performance. 


Blast is a hazard and, like smoke and the muz- 
zle flash of guns, reveals firing positions. As a 
result of developments toward the end of World 
War II, temperature effects are now much less 
restrictive. 

Largely because of the properties enumerated, 
the principal tactical uses for which rockets have 
been preferred over shells and bombs are the fol- 
lowing: 

1 . Firing heavy projectiles from shoulder launch- 
ers, small surface craft, light vehicles, and, perhaps 
most important, airplanes. 

2. Drenching area targets with intense barrages 
for short, though usually critical, periods. 

3. Firing from ground locations to which trans- 
portation of guns capable of comparable effects is 
difficult or impossible. 

4. Attacking underwater targets like submarines 
or ship hulls and underground targets like caves. 

The tactical situation in view will usually indicate 
roughly the specifications to be met as to range, 
velocity, dispersion, weight of payload, total weight, 
fuzing, and type of launcher. In general, it is de- 
sirable to provide launchers to fit the final rocket 
design, but frequently considerations of launchers 
already available or of available sizes of tubing 
from which to make rockets limit the choice of 
calibers. The following sections show in general 
terms what combinations of some of these factors 
can be met with conventional solid fuel rockets. 
Later chapters cover rocket principles, design, and 
performance in greater detail. 

In addition to their uses as parts of projectiles, 
solid fuel rocket motors have found employment 
as thrust units for assisting the take-off of airplanes 
and of long-range jet-propelled missiles, with and 
without wings, for propelling oversize fins through 
the air as targets for antiaircraft gunners, and for 
projecting lines, cables, and nets for clearance of 
land mines, and for other similar uses. However, this 
and the following chapters will be concerned only 
with rocket projectiles, and mainly with those types 
developed during the years 1941 to 1945 at the 
California Institute of Technology under Division 3, 


117 


118 


MILITARY NEEDS WHICH ROCKETS CAN MEET 


Contract OEMsr-418. Most of these rockets were 
adopted by the Army or Navy, or both. Ballistite, 
the double-base composition used in trench mortars, 
was the propellant used in all of these. With the 
exception of “Tiny Tim,” the 12-in., 1,200-lb air- 
craft rocket, all of them used single-grain charges. 


ranges by better streamlining. Another requirement 
for very long ranges is the extension of the propul- 
sion phase, that is, of the burning time of the 
propellant. A fuller discussion of the range problem 
is given in reference 1 . 

A more practical question than that of the ulti- 



0 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000 11000 12000 

RANGE IN YD 

Figure 1 . Payload vs range. 


142 RANGE 

The maximum range of the rockets considered 
here is not much greater than 10,000 yd. Attain- 
ment of very long range in ground firing is primarily 
a matter of minimizing supersonic air drag and 
secondarily one of maximizing the velocity at the 
end of the propulsion phase, since this phase is a 
small part of the trajectory length. This point is 
expanded in Section 21.2. Figure 4 of Chapter 21 
shows the effects of air drag and initial velocity on 
range. CIT put little effort into attempts to extend 


mate range is that of the range variation with pay- 
load. Figure 1, taken from reference 2, summarizes 
the data on this point, comparing service rockets 
with fixed and semifixed shells for howitzers. It is 
apparent from the figure why rockets have not been 
used to a significant extent for ground or sea firing 
at ranges beyond 5,000 yd. 

143 VELOCITY AND PAYLOAD 

For a fixed weight of payload (head), the velocity, 
and hence the range, of a fin-stabilized rocket can 


VELOCITY AND PAYLOAD 


119 


vary between wide limits, depending on the size 
of the motor or, ultimately, on the amount of 
propellant in the motor. Chapter 22 discusses 
briefly the limits on the amount of propellant which 
can be put into a fin-stabilized rocket motor of a 
given diameter. Theoretically, the problem of at- 
taining maximum velocity is slightly different from 
that of attaining maximum propellant weight a 
because, as is apparent qualitatively from Figure 12 
of Chapter 22, the use of a thicker web than that 
corresponding to the heaviest possible grain allows 


GRAIN WEIGHT IN POUNDS 



Figure 2. Maximum payload and velocity for a 
series of 5.0-in. high-capacity spinners with cruci- 
form grains. 

a considerable reduction in motor length, with a 
consequent weight reduction which more than com- 
pensates for the decreased propellant charge. In 
practice, however, when factors of propellant 
strength as well as geometry are considered, the 
shorter, thicker grains turn out to be preferable 
even from the standpoint of maximum loading 
density. Hence, for fin-stabilized rockets, once the 
maximum grain weight has been determined, the 
velocity attainable with a motor of a given caliber 
depends only on the total weight of the rocket, being 
in fact inversely proportional to it. One can attach 
to the motor a payload as large as he likes if he 
accepts the inevitable reductions in velocity and 
range . The highest velocity so far achieved in a fin- 


a A fuller discussion of this point is contained in reference 3, 
which gives curves for determining graphically the grain 
configuration which will give maximum velocity for any motor 
weight and payload. 


stabilized service rocket is .the 1,360 fps of the 5.0- 
in. high-velocity aircraft rocket [HVAR]; this carries 
a 48-lb head. 

With spin-stabilized rockets there is much less 
freedom in the choice of payloads and velocities, 
because this type of stabilization imposes rather 
rigid restrictions on the ratio of length to caliber. 
With few exceptions, heads and motors of spinners 
have been of approximately equal diameters. The 
relationships between velocity and payload are well 
illustrated in the family of 5.0-in. high-capacity 
spinners [HCSR] developed by CIT. All members 
of the family have the same diameter, length, and 
total round weight. Increases in weight and length 
of the head are associated with corresponding de- 
creases in the motor and in the velocity, as shown in 
Figure 2. This illustrates the stringent restrictions 
on possible spinner performance. Thus a high- 
capacity spinner with the velocity of the fin-sta- 



Figure 3. Payloads and velocities for 5.0-in. spin- 
ners and finners. 


bilized HVAR (1,360 fps) would have a payload of 
less than 20 lb, and to match the H VAR’s 48-lb 
payload is not possible at any velocity. The spinner 
could, of course, do a little better with a payload of 
higher average density. The comparison is shown 
in a different way in Figure 3 which assumes that 
24 lb is the maximum amount of propellant which 


120 


MILITARY NEEDS WHICH ROCKETS CAN MEET 


can be put into a 5.0-in. spinner motor. The curves 
show that it is only for small payloads that spinners 
are useful, but their variation of velocity with pay- 
load is so steep that in the limit of very small pay- 
load they surpass the finners which have more excess 
weight to carry. 


ing, as a function of payload, the velocity which one 
could reasonably expect to attain with rocket mo- 
tors of various diameters, but this has not been 
possible so far because not enough information is 
available on the variation in motor weight with 
diameter. The nearest approach that can be made 



Figure 4. Rocket velocity vs payload. With the two exceptions noted, all CIT service rockets fall within 
the triangular area included inside the dashed line. 


If it is desired to impart a certain velocity to a 
certain payload irrespective of the motor diameter, 
this can, of course, be done with either a spinner or a 
finner, but the latter can have a smaller diameter. 
Because of the marked increase in fabrication dif- 
ficulty with increasing diameter, particularly for the 
propellant, one will not choose the spinner unless it 
has distinct advantages from some other point of 
view. 

It would be useful to compile a set of curves show- 


is shown in Figure 4, which illustrates clearly the 
importance of minimizing motor weight if high 
velocities are desired. Thus, regardless of how small 
the payload is, no rocket of any diameter can 
exceed 2,800 fps unless its ratio of motor weight 
(including propellant) to propellant weight can be 
brought below 2.0. This ratio decreases with in- 
creasing diameter; for service finners its value 
averages 5.0 for 2.25-in. motors, 3.0 for 3.25-in. 
motors, and 2.7 for 5.0-in. motors, but extrapolation 


CHOICE OF FIN OR SPIN STABILIZATION 


121 


beyond this point is extremely uncertain. 15 One 
spinner (the 5.0-in. Rocket Mk 7 Mod 0) actually 
reached 2.0 and has the highest velocity of any CIT 
service rocket — 1,540 fps. With two exceptions, all 
CIT rockets fall within the narrow triangular area 
marked off in Figure 4. The exceptions are the 2.25- 
in. subcaliber aircraft rocket [SCAR], which is not 
strictly comparable with the others because it car- 
ries no payload, and the 5.0-in. Motor CIT Model 
38 (assumed to have the same payload as the 
HVAR), which was deliberately designed to have 
the lightest possible motor by accepting a lower 
safety factor (narrower temperature limits) than 
that of the 5.0-in. HVAR and other service rockets. 
It is important to note that the ratio, motor weight 
to propellant weight, is directly proportional to the 
test pressure (i.e., operating pressure times safety 
factor) and inversely proportional to the tubing 
tensile strength for any caliber of motor, neglecting 
heating effects. 28, (See Chapter 23.) Hence it is not 
possible to design an efficient rocket falling far out- 
side the triangular area in Figure 4 unless one 
employs lower safety factors, lower operating pres- 
sures, or higher tensile strengths than have been cus- 
tomary, or unless one goes to interior-burning 
grains so that the use of light metal alloys for motor 
tubes is possible . (See Section 23.2.6.) 

144 ACCURACY 

The factors determining a rocket’s dispersion are 
relatively involved and are discussed in Chapters 
24 and 25. Without attempting to indicate the 
reasons, we can summarize the dispersions attain- 
able with various types of rockets as follows: 

1 . Low-velocity ( 700 fps or less ) fin-stabilized rock- 
ets fired from typical stationary launchers. With 
burning time (duration of thrust) of approximately 
0.5 second, dispersion will be large — well above 20 
mils and perhaps above 30. It can be decreased by 
decreasing the burning time, however, and hence 

b The ratios quoted are all for motors with single-grain 
charges. The 14. 75-in. motor for the “Tiny Tim” aircraft 
rocket employed a four-grain charge. With the 18-in. extrusion 
press being completed at the Naval Ordnance Test Station, 
Inyokern, California, it will be possible to produce a single- 
grain charge for a motor of this size. With conservative 
design, an octoform grain of probably 175 lb could be accom- 
modated. With the present charge support eliminated and 
with the use of lightweight fins, the loaded motor would weigh 
only about 330 lb, giving a ratio, comparable to those above, 
less than 1.9. Still lighter motors may be practicable. 


will vary markedly with temperature. If burning 
times are brought down to 0.1 or 0.2 second as by 
use of thin- web grains, dispersions less than 10 mils 
are attainable. If all the burning can be made to 
take place on the launcher, 0 the dispersion will, 
of course, be only 2 or 3 mils. The short burning 
times are feasible only with small payloads or small 
velocities if single-grain charges are used. d 

2. High-velocity ( 700 to 1,400 fps) fin-stabilized 
rockets fired from typical stationary launchers. The 
smallest dispersion obtained up to the present with 
conventional designs is just under 20 mils. No 
means are now apparent for improving this in 
service rockets. This dispersion is lower than that 
of comparable slower rockets primarily because of 
the greater length of the faster rounds. Longer 
burning times are usually required for the higher 
velocities, but at these velocities changes in the 
burning time have little effect on dispersion. 

3. Ground-fired spinners. Spinners to be fired at 
high quadrant elevations at ground targets must 
have relatively low stability in order to follow the 
curved trajectory; they have a minimum dispersion 
of slightly under 10 mils and frequently average 
almost 20 mils at high angles. Five mils or less is 
attainable with high-spin rockets which are restricted 
to flat trajectories, 4 but only with extreme care in 
manufacturing the parts. 

4. Forward-fired aircraft rockets. Fin-stabilized 
rockets have ammunition dispersions (exclusive of 
dispersion due to pilot, plane, wind, and sight) of 2 
to 5 mils, with the lower values corresponding to 
higher aircraft speeds. Spin-stabilized aircraft 
rockets had not been tested very extensively before 
the end of World War II, but dispersions of approx- 
imately 5 mils laterally and 2.5 mils vertically were 
being obtained.® 

An indication of the relative accuracy of rockets 
and shells is given by Figure 5, taken from reference 
2, in which a fuller discussion is contained. 

14 5 CHOICE OF FIN OR 

SPIN STABILIZATION 

A first and basic decision which must be made in 
designing a rocket concerns its type of stabilization. 

0 As in the bazooka. 

d With multiple-grain charges, larger loads can be given 
higher velocities, but only by accepting higher motor weights. 

e This development was continued under the Bureau of 
Ordnance. 


122 


MILITARY NEEDS WHICH ROCKETS CAN MEET 


The relative advantages of the two types can be 
summarized as follows : f 

1. Simplicity and cheapness . A given impulse can 
be obtained with a fin-stabilized motor having a 
considerably smaller diameter than the necessary 
spinner motor. Because slim grains are cheaper 
than fat ones, small tubes more easily machined 
than large ones, and single nozzles cheaper than 


particular, they are easily adaptable to automatic 
launching, as finners are not unless the velocity is so 
low that a motor of diameter approximately half 
that of the head or less can be used with a ring tail 
(e.g., 4.5-in. barrage rocket and 7.2-in. antisubma- 
rine “Mousetrap” rocket). 

5. Aircraft armament. For firing forward, finners 
seem to be slightly more accurate, they can carry 



canted multiple nozzles, almost any rocket job can 
be done more cheaply by a finner than by a spinner. 
Also, launchers for finners are usually lighter and 
less complicated. 

2. Payload. For a given diameter and velocity, a 
finner can carry considerably more payload than a 
spinner because of the absence of a length restric- 
tion. Hence, if the caliber is fixed, there are many 
rocket jobs which cannot be done by spinners at all. 

3. Accuracy. Except in the limited region where 
very short burning times can be used, greater 
accuracy is attainable with spinners. 

4. Handling. Their stubbiness and lack of pro- 
jecting fins makes spinners more easily handled. In 


larger payloads, and they can be fired from simple, 
external, low-drag launchers. Spinners, but not 
finners, are readily adaptable to firing from within 
the wings or fuselage. For firing in other than the 
forward direction, only spinners offer possibilities. 

6. Underwater stability. Finners can be made 
stable for considerable lengths of underwater or 
underground trajectory, whereas spinners probably 
cannot. (See, however, Section 25.9.) 

7. Versatility . Spinner heads and motors must be 
matched to each other for each application, one type 
of round for aircraft use, another for accurate, flat- 
trajectory ground fire, and a third type for high- 
angle fire, necessarily less accurate. A single finner 
motor, on the other hand, can be used with many 


1 See also reference 5. 


EFFICIENCY OF ROCKET ARTILLERY 


123 


heads for many purposes; in ground fire any of the 
resulting rounds can be used at all angles of eleva- 
tion. Most commonly, finners are employed at high 
angles, with low accuracy, for area barrages. 


146 EFFICIENCY OF 

ROCKET ARTILLERY 

Questions have been raised frequently as to the 
efficiency of rockets as compared to other forms of 
artillery. These questions are applicable, of course, 
only in those situations in which it is possible to 
achieve the desired effects at the target with at least 
one of the other forms of artillery — field guns, ma- 
chine guns, aircraft bombs, aircraft cannon, etc. — 
and only when the alternate form of artillery can be 
made available in the necessary quantity at the 
necessary time and place. 

The efficiency of artillery can be evaluated in 
various ways. Rockets can be compared (idealisti- 
cally) with guns in terms of “thermodynamic effi- 
ciency/ ' measured by the ratio of the kinetic energy 
acquired by the projectile to the total energy re- 
leased by the burning of the propellant. Overlook- 
ing heat losses, this reduces to “propulsion effi- 
ciency.” A simple comparison is that between the 
amounts of the same propellant needed in a gun and 
in a rocket to give the same velocity to projectiles 
of equal masses. On this basis, rockets are con- 
siderably less efficient than guns — for example, to 
give a 25-lb rocket a velocity of 700 fps, 2.5 lb of 
propellant are required, more than ten times the 
amount needed in a mortar to fire a shell of about 
the same weight at this velocity. 

An explanation is provided by the principles of 
mechanics. In each case the energy available from 
the powder is divided between the projectile and a 
second agency in such a way that the momentum 
(product of mass and velocity) of the projectile is 
equal to and opposite to that of the second agency. 
With a few simplifying restrictions, we can make the 
following analysis of gun and rocket action: 

Mi = mass of the projectile. 

M2 = mass of second agency. 

V i = velocity of projectile. 

Vi = velocity of second agency (in free recoil). 

MiVi = momentum of projectile. 

MiVi = momentum of second agency. 

Ei = y&MiVi 2 = energy absorbed by projectile. 

Ei = y 2 MiVi* = energy absorbed by second agency. 


From the law of conservation of momentum, 


MiVi 

= M 2 v 2 , 

(1) 

from which 



v 2 

MiVi 

m 2 

(2) 

From the preceding definitions, 


E 2 

= Hmw, 

(3) 

so, from equation (2), 



e 2 

_ (M1F1) 2 

c M 2 y ’ 

(4) 

which reduces to 



E 2 

/2 ' 11 (Mt)’ 

(5) 


and, from the definition of Ei, 


E 2 — E i 


(Mi) . 

(M,) 


( 6 ) 


In a gun, the second agency includes all the re- 
coiling components. As the last equation indicates, 
the energy absorbed by these is less than that given 
the projectile by the ratio Mi/M 2 of the mass of the 
projectile to the much larger mass recoiling. Thus 
most of the energy goes into the projectile. In a 
rocket, on the other hand, the second agency is the 
propellant gas ejected at high velocity, and the 
energy this absorbs is more than that given to the 
projectile by the ratio M\/M 2 of the projectile 
mass to the much smaller mass of propellant. For 
service rockets the ratio Mi/M 2 varies from 5 to 
40, that is, the projectile may receive as little as 
V 40 of the energy available from the propellant. 

In the preceding analysis, the gun suffers by the 
assumption (true for rockets) of free recoil. The 
effect of restraining the recoil of a gun is to increase 
the “efficiency” beyond that indicated in the pre- 
vious paragraph. The amount of propellant re- 
quired in a gun is proportional to the square of the 
projectile velocity; in a rocket it is proportional to 
the first power. Consequently, the apparent “effi- 
ciency” advantage of the gun becomes less spec- 
tacular at higher velocities. However, as shown in 
Figure 6, it is maintained well beyond the velocities 
obtainable with service rockets, with their rela- 
tively short burning time. Another factor not cov- 
ered by “thermodynamic efficiency” becomes more 
important at the higher velocities; the percentage of 
payload in the rocket becomes less. 



124 


MILITARY NEEDS WHICH ROCKETS CAN MEET 


APPROXIMATE MAXIMUM RANGE IN YARDS 

250 1000 2000 4000 8000 10,000 12,000 16,000 20,000 




Figure 7. Payload vs total weight of equipment. 


EFFICIENCY OF ROCKET ARTILLERY 


125 


More important than efficiency in the use of the 
energy available from the propellant, from the prac- 
tical standpoint, is the concept of “military effi- 
ciency .” This involves a comparison of the amounts 
of effort required to inflict specified damage on the 
enemy by various means. This is, of course, an 
extremely complex problem, but one factor in it 
can be evaluated by considering the ratio of pay- 
load delivered to the target to the total weight of 
material which is to be transported to the firing 
point to deliver that payload. In the matter of 
weight, the rocket has a great advantage because its 
launcher is so light. The weights of the standard 
launchers (most of them automatic or multiple) 
for fin-stabilized ground-fired rockets range from 7 
to 37 lb per round. Although an exact comparison 
with guns involves questions of rate and amount of 
fire required, the advantage obviously lies with the 
rocket. 

On the other hand, the rocket suffers from the 
disadvantage that it must carry along its motor, 


which is usually dead weight from the standpoint of 
usefulness at the target. This handicap increases 
with velocity. Hence the velocity or range required 
affects the choice between rockets and guns as to 
whether a given amount of payload can be de- 
livered to the enemy with a smaller total weight of 
equipment. An analysis 2 based on the average 
weights of various kinds of equipment yields the 
graphs shown in Figure 7. At the points of inter- 
section (which are marked) between the shell curve 
and a rocket curve for a particular range, the total 
amount of equipment necessary to lay a given quan- 
tity of effective ammunition (payload) on the 
target will be the same for both rockets and shells. 
Below these points, rocket propulsion will be more 
“efficient.” Evidently it is at short ranges that 
rockets have the most distinct advantage, in con- 
trast to the situation for thermodynamic efficiency. 
During World War II, large numbers of rockets 
were used for area barrages at ranges from 1,000 
to 5,000 yd. 



Chapter 15 

ROCKET HEADS 

By C. W. Snyder 


151 SIMILARITY TO 

SHELLS AND BOMBS 

R ocket heads, exclusive of their fuzes, have 
been the subject of relatively little experi- 
mental investigation. In many cases they have 
been adapted with relatively minor modifications 
from standard shells or bombs, which is reasonable 
since they are intended to do substantially the same 
job at the target. From the point of view of per- 
formance at the target, the problems of exterior 
contour, optimum well thickness, steel composition 
and heat treatment, etc., for rocket heads are 
generally similar to those for the corresponding 
shells or bombs. 

152 ALIGNMENT 

The relation of the head to the motor does present 
certain unique problems, of which the foremost is 
the matter of alignment. The meticulous care which 
is taken to assure proper alignment of nozzle axis 
and motor tube to ensure low dispersion is obviously 
of no avail if the center of mass of the head is far 
from its axis, so that comparable precautions must 
be taken in head manufacture. Well thicknesses 
must be relatively uniform, filling must be sym- 
metrical, and threads for attaching to the motor 
must be machined so that their axis passes through 
the center of the mass of the head within the re- 
quired accuracy. It has been customary to use the 
same thread specifications on heads as on motors 
(see under Alignment in Section 23.2) although 
obviously the precision required for head threads 
depends markedly on the length of the head and its 
weight relative to the total rocket weight. In any 
particular case, it is necessary to calculate the effect 
that various types of head malalignment have on 
the overall round malalignment and adjust toler- 
ances accordingly. If the head is the major portion 
of the rocket weight, it may be desirable to balance 


each one. a For fin-stabilized rockets, the goal is to 
keep the possible overall malalignment of the round 
under about Vio degree. Although the limit for 
spinners is not established there are clear indications 
that dispersions as low as 5 mils (mean deviation) 
are unattainable unless each main component, and 
preferably also the assembled round, is dynamically 
balanced. 

153 LEAKAGE AND HEATING 

The base of the head serves usually as the front 
closure of the motor chamber; its exposure to the hot 
gas in the motor creates problems in some cases. 
Thus, at one time, concern was felt about the heat- 
ing of the TNT in the head until tests showed that 
for the short burning times being used 34 in- of steel 
was sufficient insulation. Inferior steel bar stock 
may sometimes contain longitudinal “pinholes,” 
however, so that a certain amount of care is still 
required in manufacturing the base portions of heads 
and the connectors between rocket motors and heads. 

Gas leakage around base fuzes presents a similar 
problem, which is discussed in greater detail in 
Chapter 16. 

15 4 JOINT STRENGTH 

In cases where the motor is required to remain 
attached to the head after impact, the strength of 
the joint between the two becomes critical. For this 
reason the underwater heads for the 3.5-in. aircraft 
rockets have long “skirts” which extend back of the 
threads, and the 5.0-in. high-velocity aircraft rocket 
heads have their connecting threads 3.5 in. forward 
of the base (see Figure 1). A similar construction 


a Methods for balancing the heads of fin-stabilized rockets 
are given in reference 1. Spinner heads require dynamic bal- 
ancing; equipment for this is discussed in a CIT final report on 
testing of rockets. 2 


J26 


SPECIAL HEAD SHAPES 


127 


was originally adopted for “Tiny Tim,” the 11.75- 
in. aircraft rocket, but was abandoned for various 
minor reasons after tests showed that it was not 
required to prevent breakup on water impact. It 
may be that it is necessary, however, to minimize 
the frequency of breakups on ground impacts, so 
that a redesign will be required if the potentially 
long underground trajectory of this rocket is to be 
utilized (see Chapter 24) . 


4'.'0 



3l'260 +.010 



B 

Figure 1. Reinforced motor-head connections for 
aircraft rockets, (A) HVAR, (B) 3.5-in. AR. Di- 
ameters of critical surfaces are shown. 

SPECIAL HEAD SHAPES 

The shape of a rocket head is usually important 
mainly from the standpoint of effectiveness at the 
target, i.e., achieving maximum blast effect or most 
efficient fragmentation. Occasionally, however, the 
shape may affect the trajectory to the target. The 
first case of this kind was encountered with the 
antisubmarine rocket [ASR], which is shown in 
Figure 1 of Chapter 18. The flat nose of its head 
was originally copied from the British “Hedgehog” 
projectile, b and extensive underwater trajectory 
tests at CIT soon demonstrated that it was superior 
to various other nose shapes suggested because its 
use resulted in smaller forward travel after impact 


and less deviation from the mean trajectory. The 
reason apparently is that its very large drag causes 
it to be decelerated to less than its terminal velocity 
during the first 10 ft of underwater travel, after 
which it sinks almost vertically with increasing 
speed. The underwater behavior of the ASR and of 
its cousins the VAR’s c with various head shapes, 
tail shapes, fuzes, etc., are discussed in many re- 
ports by the CIT Morris Dam group. d 

The control of the underwater trajectory of air- 
craft rockets is also a matter of head shape. The 
fact that fin-stabilized rockets fired forward from 
aircraft have long, accurate underwater trajectories 
was discovered by the British, and extensive tests 4 
by CIT showed that it was possible by proper atten- 
tion to head shape to increase the effective under- 
water range considerably and to introduce a certain 
amount of control over the curvature of the rocket’s 
path. It is well known that a rapidly moving pro- 
jectile under water moves in a bubble as illustrated 
in Figure 2. The water is, of course, held in direct 



Figure 2. Position of rocket under water. 


contact with the nose of the projectile, but at some 
point ahead of the cylindrical portion of the projec- 
tile the water recedes from the axis faster than the 
ogival radius of the projectile increases, so that, in 
the absence of gravity, the water would touch the 
projectile nowhere except at the nose. Actually, 
the rear of the projectile drops to the bottom of the 
bubble and rides in the water deep enough so that 
the force of the water on it balances the projectile’s 
weight. Under these circumstances, the resisting 
force experienced by the projectile depends upon 
the energy imparted to the water or, in other words, 
entirely upon the diameter of the bubble and not at 
all upon the diameter of the projectile. The diameter 
of the bubble, and hence the resisting force, can be 
reduced by means of the so-called “double-ogive” 

c Vertical antisubmarine rockets, also known as retro rockets 
and retro bombs. 

d The work of this group is summarized in a CIT final 
report. 3 In Chapter 1 of this volume Max Mason gives an 
introductory survey of this work. Reports on it are listed in 
the CIT OEMsr-418 bibliography in the general bibliography 
in the appendix. 


b A small spigot-projected antisubmarine depth bomb. 



128 


ROCKET HEADS 


head; this has a small radius of curvature near the 
tip of the nose, blending into a curve of much larger 
radius which joins the straight section at the rear of 
the head. 

Since the rocket travels in its bubble with usually 
an up yaw, the reaction of the water on the nose is 
not, in general, symmetrical. An upward force exists 
which depends greatly on the shape of the ogive 
at the tip. A hemispherical ogive, since it presents 
the same appearance to the water even when rotated 
at a small angle, has almost no upward force. As 
the ogive is made sharper, the upward force in- 
creases. Hence, within the limits of force which the 
rocket can stand without breaking, one can obtain 
almost any value of upward force and hence control 
the curvature of the trajectory by changing the 
sharpness of the nose. 

Three typical heads for the 3.5-in. aircraft rocket, 
and their performance, are shown in Figure 3. Vari- 
ous other head shapes for 2.25-in., 3.25-in., 5.0-in., 
and 11.75-in. aircraft rockets are discussed in re- 
ports issued by CIT under Contract OEMsr-418. e 

A rocket penetrates earth or concrete in a manner 
essentially identical with that in which it penetrates 
water, so that the theory of head shapes should be 
the same. This is, in fact, found to be the case, 
except that the forces in solids are much greater 
than in liquids so that restrictions on possible head 
shapes are tighter. Also, as mentioned previously, 
the strength of the joint between motor and head is 
much more critical. The underground trajectories 
of aircraft rockets with various heads are discussed 
more fully in Chapter 24. 

e See the CIT OEMsr-418 bibliography in the general 
bibliography in the appendix. 



0?* 7( 

)" 




B 



Figure 3. 3.5-in. underwater heads. 


A. Single-ogive Mk 1 
Deceleration coefficient 0.0136 
Radius of curvature 200 ft 
Distance to half velocity 51 ft 

B. Sphere-ogive 
Deceleration coefficient 0.0065 
Radius of curvature infinite 
Distance to half velocity 107 ft 

C. Double-ogive Mk 8 
Deceleration coefficient 0.0069 
Radius of curvature 620 ft 
Distance to half velocity 100 ft 


Chapter 16 

ROCKET FUZES 


By C. W . Snyder 


161 GENERAL REQUIREMENTS 

R ocket fuzes, like those for bombs and pro- 
» jectiles, have two prime functions: (1) to dis- 
perse, ignite, or, usually, detonate the contents of 
the rocket head under the proper circumstances, 
and (2) to prevent such actions under all other con- 
ceivable circumstances. Because these two basic 
requirements are distinct, a fuze mechanism can 
usually be thought of functionally as two sets of 
interrelated mechanism: (1) the firing mechanism , 
which performs the end functions, and (2) the arm- 
ing mechanism , which prevents firing until com- 
pletion of a sequence of operations which depend 
on some of the phenomena associated with the 
launching and flight of the rocket. Arming is com- 
pleted when all of the elements in the explosive 
train, loosely called the “detonator,” are uncovered 
and in line with the firing pin, ready to function on 
impact or some other stimulus. 

Fuze design is a specialized business, consisting 
mostly of modifications of a relatively few basic 
types so that they are usable with rockets with 
drastically different characteristics of pressure, ac- 
celeration, burning time, and tactical use. It can 
only be summarized here, mainly from the more 
complete discussion of wartime fuze work at CIT 
given in a Rocket Fuzes. 1 The following discussion 
shows how fuze problems may affect the design of 
other rocket components and indicates the general 
types of fuzes worked on at CIT. b With one excep- 
tion, all these fuzes are mechanical and differ from 
standard bomb and projectile fuzes mainly in their 
methods of arming. With the same exception (the 
fuze for ejection of “window”), firing of all of these 
fuzes is accomplished by percussion, by the im- 
pinging of a firing pin on a pellet containing a small 
quantity of sensitive explosive. 

a One of the final report volumes issued by CIT under 
Contract OEMsr-418. 

b For information on other rocket fuze developments in 
NDRC, see (1) references 2 and 3; (2) Division 8 Summary 
Technical Report on fuzing of shaped-charge heads; (3) Divi- 
sion 4 Summary Technical Report on proximity fuzes for 
rockets. 


The safety requirements for rocket fuzes are sub- 
stantially the same as those for other projectile 
fuzes: the arming system should provide restraints 
on the firing mechanism, and these restraints should 
remain effective under the forces of transportation, 
handling, loading, and launching. Many of the 
rocket fuzes developed during World War II do not 
entirely satisfy the usual safety requirements. 

These requirements are usually more difficult to 
meet in rocket fuzes than in projectile fuzes, because 
of the smaller margins between the forces imposed 
by careless handling and those available for actua- 
tion of the arming mechanism. For this reason it is 
frequently necessary to utilize for arming a com- 
bination of forces such that the probability of their 
simultaneous occurrence under circumstances other 
than launching and flight of the rocket is negligibly 
small. In most rocket fuzes, as in projectile fuzes, 
arming is made to depend on phenomena associated 
with launching and flight of the projectiles in which 
they are mounted, and is completed only after a 
period of projectile flight. 

162 METHODS OF ARMING 

The initiation of the arming process in many fuzes 
for fin-stabilized rockets, especially those fired from 
aircraft, depends on withdrawal of a wire, similar to 
the arming wire used on bomb fuzes. Among the 
arming methods not dependent on the conditions of 
rocket launching and initial flight are water pres- 
sure, spring-driven flywheels, and deceleration 
changes. These methods are used only for special 
target situations. Most rocket fuzes depend for 
arming actuation on one or more of the following 
conditions associated with projection; note that two 
of these conditions are peculiar to rockets . 

Acceleration Forces (Setback). In guns, the accel- 
eration of the projectile is very large (14,000*7, c for 
example, in 5.0-in. naval guns) so that setback can 

c g = 32.2 ft/sec 2 = acceleration of gravity at the surface 
of the earth. In this example, each element of the projectile 
is accelerated by a force 14,000 times its weight. 


129 


130 


ROCKET FUZES 


readily be used as a primary arming force. In 
rockets, on the other hand, acceleration is not only 
small, being never more than 100# in CIT rockets 
and in extreme cases falling as low as 3 g, but also it 
varies over a wide range with the propellant tem- 
perature. Hence setback cannot safely be used as a 
primary arming force except in conjunction with 
other forces. It is often used to delay completion of 
arming until the end of burning. 

Wind Forces. The force provided by the wind 
streaming past the rocket in flight is most frequently 
utilized to arm nose fuzes (usually by a propeller) 
on finned rockets. On rockets having supersonic 
velocities the air pressure at the nose can be used 
for arming nose fuzes. 

Pressure of the Propellant Gas. Since the pressure 
in a rocket motor is relatively large, it can be used 
conveniently to arm base fuzes. The burning times 
are usually long enough that the entrance of gas 
into the fuze can be delayed so that arming does not 
begin until the rocket is well beyond the launcher. 

Heat of the Propellant Gas. The hot gas can be 
used to initiate delay powder trains of special types . 

Centrifugal Force. Most shell fuzes make use of 
the large radial forces set up by the spin, and, since 
spinning rockets have rates of spin comparable to 
those of shells, similar or, in some cases, identical 
fuzes can be used for them. In fin-stabilized rockets, 
of course, these forces are absent. 

The various types of fuzes developed by CIT 
during World War II were designated by three let- 
ters, the first to indicate the method of arming, the 
second the method of firing (I for impact) and the 
third the type of projectile (R for rocket in all 
cases). The following list shows type designations 
(roughly in the order in which developments started) , 
modes of arming, and the Mark numbers assigned 
by the Navy Bureau of Ordnance to specific fuzes 
of these types: 

HIR, armed by Hydrostatic pressure, Mks 135 
and 140. 

AIR, armed by rotation of a propeller in the 
Airstream, Mks 137, 147, 148, 149. 

SIR, armed by a Spring-rotated shaft, Mk 139. 

NIR, armed by air pressure on the Nose, Mk 144 
(never standardized or used) . 

PIR, armed by Pressure of the propellant gases 
in the motor, Mks 146, 157, 159, 163, 164, 165. 

DDR, a special type of the PIR in which firing 
depends on Deceleration Discrimination. 


163 AIR NOSE FUZES 

All nose fuzes which have had extensive use on 
CIT fin-stabilized rockets have been modifications 
of the AIR fuze, one of which is shown in Figure 1. 
At least twelve modifications have been developed 




CAP SPRING 
PROPELLER HUB RIVET 
GASKET 

CLAMP 
NOSE PLATE 

PROPELLER LOCKING 
PELLET 

SET-BACK SPRING 
BODY 

FIRING PIN GUIDE 


LEAD-IN (TETRYL) 
BOOSTER (TETRYL) 


Figure 1. Mk 149 (AIR) nose fuze. 


for the 4.5-in. barrage rocket, the 7.2-in. chemical 
warfare rocket, and all the aircraft rockets from 3.5 
in. up. In all of them, (1) acceleration retracts a 
setback block, thus unlocking a propeller which, 
driven by the airstream, turns a shaft in threads to 
free the firing pin and complete the arming, and (2) 


NIR NOSE FUZES 


131 


firing, instantaneous in most cases, is by percussion 
on impact with the target; the fuzes are point- 
detonating . 

The arming characteristics of AIR fuzes are fitted 
to the various rockets primarily by varying the size 
and shape of the propeller and the pitch of its blades . 
For use on aircraft rockets at high velocity, the 
Mk 149 fuze (AIR 8) has a streamlined body, and 
it is protected from corrosion or from the possibility 
of being fouled with ice by enclosing the propeller 
in a metal cap which is thrown off by a spring on 
removal of an arming wire at the time of firing. For 
use on rocket heads shaped for good underwater 
trajectory, the AIR 12 has a roughly hemispherical 
body. The AIR 9, 10, and 11 had special propellers 
and delay detonators to make them “water-dis- 
criminating,” to fire on or after underwater hits 
against ships. Because of the difficulty of keeping 
a nose fuze intact long enough for it to operate with 
a delay on armor plate of appreciable thickness, 
work on these models was finally abandoned in 
favor of the DDR base fuze, described in Sec- 
tion 16.6. 

164 NIR NOSE FUZES 

Fairly extensive experiments were made with a 
drastically different type of point-detonating nose 
fuze, the NIR. Air, compressed in front of the 
rocket, enters through ports at the nose into a cham- 
ber around a sylphon bellows (see Figure 2) , which 
collapses and retracts the firing pin, thus allowing 
the detonator shutter to move into place and com- 
plete the arming. Development work on this fuze 
was not completed because it was found that, at 
subsonic velocities, the nose pressure was not 
enough greater than atmospheric to allow the fuze 
to function with sufficient reliability. For super- 
sonic velocity, however, the NIR should be reliable, 
and it is suggested in Rocket Fuzes 1 that its use 
would have distinct advantages in the following 
cases: 

1. For high-velocity aircraft rockets, since the 
partial arming distance could be increased above 
that obtainable with AIR fuzes; 

2. For slow-spin fin-stabilized rockets with a high 
velocity, since the spin is great enough to prevent 
the proper functioning of a propeller-arming fuze 
but not sufficient to allow arming by centrifugal 
force; 


3. For spinners which attain their full spin 
velocity on the launcher; and 

4. For antisubmarine rockets (arming by hydro- 
dynamic pressure). 




Figure 2. Mk 144 (NIR) nose fuze. 


16 5 PIR base fuzes 


Mechanical base fuzes are used in preference to 
nose fuzes chiefly when it is desired that the head 
detonate after impact with a certain delay so that 


132 


ROCKET FUZES 


it will penetrate armor instead of blowing up outside 
it. Base fuzes are fired by the inertia of certain of 
their parts which tend to continue in motion when 
the head is decelerated by impact. A small delay 
(up to about 0.015 second, the exact amount de- 
pending on the fuze design and the resistance of the 
target) is thus inherent in their construction, but, if 
longer delays are required, they can be achieved 
by including a pyrotechnic delay in the firing train 
(0.02-second delay has usually been used) since the 



Figure 3. Mk 146 

head itself protects the fuze from being crushed 
before the delay element can function. 

Most of the mechanical base fuzes which have 
been used in CIT fin-stabilized rockets are modifica- 
tions of the PIR, one of which is shown in Figure 3. 
Arming of these fuzes is accomplished in the fol- 
lowing manner. The motor gas, filtered free of solid 
material by the inlet screen (see Figure 3), enters 
the pressure chamber through the small orifice in 
the inlet screw, so that pressure in the chamber 
builds up slowly during burning. The pressure 
exerts a force on the diaphragm, and, when the 
force becomes large enough, it shears the shear wire 
and depresses the arming plunger, releasing the 


locking ball, which completes the first stage of 
arming. The second stage of arming is not com- 
pleted until the end of burning. 

To adapt PIR fuzes for use on different motors, 
the following modifications can be made. (1) The 
diameter of the inlet orifice can be varied so that 
the leakage of gas into the pressure chamber will be 
rapid enough to reach the necessary pressure before 
the end of burning at all temperatures expected, but 
slow enough to provide the delay required for safety. 



(PIR) base fuze. 

Diameters in the range from 0.0145 to 0.033 in. 
have been used. (2) The diameter of the shear wire 
can be varied so that the pressure necessary for 
arming is just slightly less than the motor pressure 
at the lower temperature limit. (3) The inlet screen 
and cup may be replaced with other types of filters 
and shields as may be required to keep the debris 
and the closure disks in the motor from clogging the 
tiny orifice. The first motor on which the PIR fuze 
was used (the 3.25-in. Mk 7 motor) contained a 
cellulose acetate igniter case, two fiberboard closure 
disks, and some cardboard sleeves at the front end, 
the combination of which created a considerable 
filtering problem. When this was realized, an effort 


DDR BASE FUZES 


133 


was made to clean up all motors with which base 
fuzes were to be used. With the later designs con- 
taining metal case igniters and steel closures with 
“blowout patches” (see Chapter 23), much less 
clogging has been encountered. 

16 5 1 Gas Seals 

One of the crucial problems that arises when a 
base fuze is used is that of making an effective seal 
between the fuze and the head so that the hot high- 
pressure gas from the motor cannot reach the high 
explosive either in the fuze or in the head. The 
sealing of the inside of the fuze itself is purely a fuze 
design problem and need not concern us here. For 
sealing the space between the fuze and the head, 
early PIR fuzes had a soft copper gasket such as 
that shown in Figure 3. No particular difficulty 
with leaks past the gasket had been noted with 
static firing in connection with fuze testing, but, 
when a head, in which the base fuze had appar- 
ently been omitted so that the gas had direct access 
to the TNT, detonated low order on the launcher, 
the whole problem was extensively reinvestigated. 
The results of this investigation are discussed in 
the weekly progress reports, 4,5 It was concluded 
that copper gaskets approximately 0.050 in. thick, 
annealed soft, provide adequate sealing if (1) the 
seating surfaces on the fuze flange and in the head 
are square with the threads, are smooth, clean, and 
free from defects, and are held within close toler- 
ances; (2) the gasket is in good condition; and (3) 
the fuze is screwed in with a large torque and 
seated tightly on the gasket. The tests did not show 
that prematures could result from leaks such as 
occur past a poor seal, but it was realized that this 
is a statistical matter and that even small leaks 
should not be tolerated. The primary difficulty with 
the gasket seal is that no way exists by which a bad 
assembly can be detected after it is made. 

As an alternative, gas seals of the type used in gun 
projectiles were extensively investigated. 5 a In these 
seals, a copper-encased lead “gas check” is forced 
into a triangular groove, the sides of the triangle 
being the edges of the rocket head and the fuze 
respectively (see Figure 4). Such gas checks were 
found to be entirely satisfactory if crimped in place 
with sufficient pressure, even when the parts were 
poorly assembled or had scratches or gouges on the 
seating surfaces or threads not at right angles to the 


seating surfaces. The condition of the gas check as 
seen on a visual inspection was, within limits, a 
satisfactory criterion of the effectiveness of the 
sealing. As a result of these tests, this type of gas 
check has been adopted for all base fuzes (see 
Figure 5) except those in which the fuze and the 
motor adapter are made in one piece so that no 
space for leakage exists. 




Figure 4. Gas check ring (A) un deformed and 
(B) as actually used. Center is lead and jacket is 
copper. Illustrations are about 12 times actual 
size. 

16 6 DDR BASE FUZES 

As is apparent from Figure 5, the DDR is a 
modification of the PIR from the standpoint of 
arming mechanism, but its method of firing is so 
unorthodox that it has been given a special designa- 
tion — the “deceleration-discriminating” fuze. It 
was designed for use with the aircraft rockets which 


134 


ROCKET FUZES 


INLET-VALVE BALL 
INLET-VALVE SPRING 
INLET-VALVE PLUG 

IN LET- VALVE PIN 


DIAPHRAGM- 


DIAPHRAGM DISK- 


LOCATING PIN 
ARMING-SLEEVE PIN- 


(45° OUT OF POSITION) 


DETONATOR CASE 


CLOSING PLUG 


BOOSTER DISK- 



BODY 

-ARMING SLEEVE 

SHEAR WIRE 


PLUNGER-PIN 
COLLAR 

-DETONATOR-PLUNGER PIN 


•STAKING PIN 




0 1 
I I I I I I I I I I I 

SCALE (IN.) 

Figure 5. Mk 166 Mod 1 (DDR) base fuze. 


STOP PIN 


LEAD-IN 
CLOSING DISK 


LEAD-IN CUP 


MAGAZINE 


BOOSTER PELLET 


ANTISUBMARINE FUZES 


135 


have stable underwater trajectories, and its opera- 
tion may be described briefly as follows. The initial 
impact (on water or target) unlocks a trigger 
mechanism which is controlled by the deceleration 
of the rocket. Nothing more happens as long as the 
deceleration remains more rapid than that which 
accompanies high-speed underwater travel. Decel- 
eration during armor penetration is, of course, much 
more rapid than this. When, after exit from the 
armor, the rocket is traveling through the less re- 
sistant air, the slow deceleration causes release of a 
spring-activated firing pin which initiates the ex- 
plosive train. Thus the fuze satisfies the basic 
requirements for the attack of heavy ships — 
whether the hit is above or below the water line, 
the fuze detonates after penetration, but does not 
detonate during impact on the water or on the 
ship. Since its functioning is independent of time 
delays and of length of underwater trajectory 
(within limits), it is effective against armor of any 
thickness which the rocket will penetrate, and it 
does not require great precision in the firing of the 
rocket. 

For use against certain land targets such as caves 
and pillboxes, the DDR fuze has special advantages, 
since, instead of detonating with a fixed delay after 
the first impact, it waits until the rocket penetrates 
the first obstacle completely or is brought to rest in 
it, thus considerably increasing the destructiveness. d 

Although the DDR fuze was developed too late 
to have any service use in World War II, some 
general remarks about its tactical use can be made. la 
Obviously a fuze of such unorthodox characteristics 
will be most effective only under very special con- 
ditions. To be useful under water or under ground, it 
must be used on a rocket which has a stable under- 
water or underground trajectory and does not break 
up; the characteristics of such rockets are discussed 
in Section 24.9. The fuze is rugged and will function 
after impact at not too great obliquity on fairly 
heavy plate, so that, if the full potentialities of the 
fuze are to be realized, the head must be equally 
rugged. Thus good results were obtained in experi- 
mental firings with the 5.0-in. Rocket Heads CIT 
Model 35 and Mk 2 Mod 2 having solid and heavy 
noses (adaptations of “special common” type pro- 
jectiles) . The only heads used during World War II, 
however, for reasons of availability, were modifica- 
tions of the 5.0-in. Mk 35 AA common shell, which 

d See reference 6 for discussion of its use in the 11.75-in. 
aircraft rocket against caves. 


has a hole in the nose and thin walls so that it 
breaks up on relatively thin plate. In such a head, 
the DDR would serve no useful purpose. 


167 ♦ ANTISUBMARINE FUZES 

Three fuzes for antisubmarine use on low-velocity 
rockets were designed by CIT: the HIR or “Hydro- 
static-arming, Impact-firing, Rocket” fuze (Mk 
135), the HIR 3 (Mk 140), and the SIR or “Spring- 
arming, Impact-firing, Rocket” fuze (Mk 139). In 
addition, extensive underwater tests were con- 
ducted on the Mk 131 and Mk 136 fuzes, which are 
two modifications of a British-designed fuze incor- 
porating underwater vane arming and inertia firing, 
and some redesign work was done on them in the 
light of the test results. 

The two HIR fuzes were very similar in principle, 
arming being effected by water pressure entering the 
fuze through ports in the nose and “popping” a 
phosphor-bronze diaphragm, which, through link- 
ages, unlocked certain restraints and aligned the 
explosive train. They were fired by the deceleration 
on impact with a solid object, which released a 
spring-loaded firing pin. Neither fuze was used 
extensively in service, since the Mk 131 was simpler 
to make, was available in quantity earlier, and 
exhibited superior performance in CIT's underwater 
tests. Since their operation did not depend on any 
of the characteristics of the rocket, they are not of 
particular interest to us here. A full discussion of 
their design and testing is contained in Rocket 
Fuzes) 1 diagrams and photographs can be found in 
references 7 and 8. Numerous CIT publications dis- 
cuss the underwater tests of these fuzes. 9-15 

The Mk 139 fuze, originally designated the SIR, 
was designed primarily for vertical bombing of sub- 
marines from low-flying aircraft. Since the rockets 
were fired rearward at a speed approximating that 
of the plane, their flight was somewhat unstable, 
and the fuzes had to be designed to arm reliably 
regardless of whether the rocket fell nose down or 
sideways. It was desired that the fuze fire on con- 
tact with the submarine hull either submerged or on 
the surface, so that water discrimination was neces- 
sary. To meet all these requirements, a coiled 
clock spring was used as the source of arming energy, 
accelerating a flywheel which gave the arming de- 
lay. Water discrimination was achieved by making 


136 


ROCKET FUZES 


it point-detonating. The fuze functioned satisfac- 
torily but saw little service use because of the de- 
cline of interest in vertical bombing. 

No detailed discussion of the fuze will be given 
here. Information on it may be found in Rocket 
Fuzes, 1 in reference 8, and in various CIT pub- 
lications. 7,16 - 19 


BASE FUZES FOR 
WINDOW ROCKETS 

To eject the payload of antiradar “window” from 
the heads of 3.5-in. rockets at the proper range and 
height, a time fuze was required, and the simplest 
such device appeared to be a powder train in the 
base of the head, initiated by the motor gases. A 
percussion-actuated dynamite-fuze ejector unit des- 
ignated the DU-5 was developed by CIT. Shown in 
Figure 6, it consists of a plastic case containing 
approximately 20 g of FFFG black powder, within 
which is coiled 4j^ in. of dynamite fuze (Bickford 
cord) sheathed in a vinyl chloride tube. The end of 
the fuze projects through a hole in the end of the 
case and is cemented in place, and a short length of 
Quickmatch, held in contact with the end of the fuze 
by a metal clip, assures easy ignition. Ignition is 
accomplished by the firing of a .32-caliber blank 
cartridge containing approximately one-fourth of 
its normal powder charge. A firing pin attached to a 
diaphragm sets off the cartridge when the pressure 
builds up in the motor. 

The tests involved in developing the DU-5 are 
discussed in detail in reference 20. The principal 
difficulties encountered were: 

1 . A design of firing pin was required which would 
not allow gas from the motor to leak into the head 
through the hole that was frequently opened in the 
cartridge when it fired. Gas leakage into the head 
reduced the ejection time because the dynamite fuze 
burned more rapidly at higher pressure, and in 
extreme cases the gas pressure blew the payload out 
of the head. 

2. Premature ejections were frequently obtained 
because the fuze burned through the side and ignited 
the black powder. This was eliminated by the 
vinyl chloride sheath. 

3. Erratic burning rates of the fuze caused by 
the building up of pressure from its own gas were 
eliminated by venting into the payload, which was 


relatively porous and provided a sufficient volume 
so that the pressure rise was small. 

Although the DU— 5 does work satisfactorily 
and was used in service, considerably better designs 




Figure 6. DU-5 delay ejector charge for window 
rocket heads. 

from the standpoint of ruggedness, compactness, 
and simplicity are possible using fuze materials hav- 
ing solid rather than gaseous products. CIT tests 
on two such units designed by the Catalyst Research 
Corporation are also discussed in reference 20. Al- 
though neither was satisfactory as it was, further 
research might remedy the defects. It may be 


FUZES FOR SPIN-STABILIZED ROCKETS 


137 


possible to initiate the fuze by the heat of the motor 
gas itself, thus considerably simplifying the design.® 

169 FUZES FOR 

SPIN-STABILIZED ROCKETS 

All fuzes mentioned previously were developed 
for fin-stabilized rockets. Development work on 
spinner fuzes was not nearly so extensive because 
the advent of spinners came fairly late in the CIT 
work and because standard projectile fuzes, nearly 
all of which are armed by centrifugal force, can be 
used with little modification. For the 5.0-in. Rocket 
Mk 7 Mod 1, the base fuze Mk 31 Mod 0 was used 
without any modification. Only very minor modi- 
fications were made to the Auxiliary Detonating 
Fuze Mk 44 Mods 1 and 2, which are used with the 

e Earlier work by the Catalyst Research Corporation with 
Section H of Division 3 on the development of gasless delay 
units for ejecting parachute flares from the heads of 3.25-in. 
rockets is covered in its final report listed in the general 
bibliography. 

In 1942 some development work was done by Section H, 
working with the Navy at the Naval Powder Factory, Indian 
Head, Maryland, on a delay-ejection device in which the 
action was initiated when the propellant gases in a rocket 
motor heated a metal tube to melt solder within it, to release 
a pin. 21 ’ 22 


nose fuzes of all but one of the spinner models de- 
veloped by CIT. Two point-detonating nose fuzes, 
the Mk 30 Mod 3 and the Mk 100 Mod 0, have been 
used on service spinners; both of them are modifica- 
tions of the Army M48 fuze. 

In adapting fuzes to various spin-stabilized rock- 
ets, the important factor is to have the arming 
occur as close to the end of burning as feasible. 
It may be necessary merely to use a spring with a 
different tension so that the arming mechanism will 
be actuated at a different spin velocity. For rockets 
fired at long range, the spin may drop to 75 per cent 
of its maximum value, so that, if the arming is 
reversible (as is the case with both the Mk 30 and 
Mk 100 nose fuzes), it must take place at less than 
75 per cent of maximum spin (corresponding to 
approximately half the burning distance) if the fuze 
is not to become unarmed again before impact. 
Since centrifugal force increases with the distance 
from the axis, a detent which has moved out and 
armed the fuze at a particular spin velocity is ex- 
erting considerably more force than before. It is 
therefore possible to arrange that the arming process 
will not reverse until the spin has dropped consider- 
ably below that at which it occurred. The factors 
involved in obtaining this “unbalanced” condition 
are discussed in Rocket Fuzes . lb 


Chapter 17 

ROCKET LAUNCHERS 

By C. W. Snyder 


171 INTRODUCTION 

I n the development of an effective rocket 
weapon, the proper design of launcher is no less 
important than that of the projectile itself. Never- 
theless, the space devoted to launchers here will be 
small because their problems are for the most part 
almost indistinct from those of the rocket itself 
and because they are discussed fully in two of the 
CIT final report volumes. 1,2 

Many types of launchers have been used, varying 
in complexity from simple cardboard tubes or 
wooden troughs to elaborate mechanisms for load- 
ing, aiming, and firing by remote control. Naturally 
many considerations enter into launcher design. 
The starting point is the tactical employment, the 
round to be used, and the platform or vehicle on 
which the launcher is to be mounted. These will 
determine the nature, length, and number of the 
guides, the nature of the mount, the electrical sys- 
tem, and the type of fire control. Consideration 
must be given to the control of the rocket blast and 
to such factors as the means of loading, protection 
from weather, and limitations on shipping volume 
and handling weight. These considerations for 
rockets fired from aircraft differ so radically from 
those for rockets fired from stationary platforms or 
surface vehicles that it proved efficient to have two 
distinct groups to handle the two types of launcher 
problems. This division will be observed in the 
following discussion. 

172 SURFACE LAUNCHERS 

In Rocket Launchers for Surface Use 1 a thorough 
discussion of the problems of launcher design for 
surface-fired finners and spinners is given, with 
complete descriptions and illustrations of all launch- 
ers which saw any service use. We shall not attempt 
to duplicate the material here. 

Launcher Types 

The basic function of a launcher is to support 
and guide the rocket in its initial motion. Three 


commonly used means are rail launchers, slot 
launchers, and tube launchers. In the first the 
rocket slides on two guide rails so spaced as to 
subtend an angle at the rocket axis of 90 to 120 
degrees. The rails are commonly made of formed 
sheet steel or small-diameter steel pipe. If the 
launcher is to be used on a moving vehicle, one or 
two upper rails may be added to hold the rocket 
down. Because of their small weight, rail launchers 
have been used widely for the low-velocity fin- 
stabilized rounds consisting of a head and ring tail 
(fins) of one diameter and a motor of smaller 
diameter . 

The guide rails are made as long as practicable, 
to increase accuracy, but seldom more than three 
times the length of the round. In some cases, 
launcher length has been combined with ease of 
handling by the use of folding or telescoping rails. 

Many aircraft rockets have lug “buttons’ ’ by 
which they are mounted on slotted launchers. The 
slot is a space of about z /% in. between two flat rails. 
A few slotted launchers have been developed for 
firing these aircraft rockets from ships. An example 
is the CIT Type 31 C (see Figure 6 of Chapter 19) 
discussed in Section 19.2.5. 

Rather long tube launchers have been used for 
certain finned rockets in which the fin diameters 
could be limited to those of the heads and for rockets 
equipped with folding fins. In these launchers the 
tubes were of the same nominal inside diameters as 
the rounds. Tube launchers have found even wider 
use for spinners, for with these the launcher length 
can be reduced almost to that of the round with little 
loss of accuracy. The short length makes weight less 
important. Most of the CIT spinner launchers were 
tubular, with clearance between tube and round 
provided by three or four internal guide rails. This 
type of launcher has given the best accuracy under 
service conditions. 

Single-guide launchers, into which only one round 
at a time can be loaded, are used for applications 
where portability is more important than rate of 
fire. For greater fire power, multiple launchers have 
been used extensively, with number of tubes or 
rails varying from 2 to 144. The launcher weight 


138 


SURFACE LAUNCHERS 


139 


per round is little different from that of the single- 
shot launcher. A considerable saving in weight and 
an enormous advantage in simplicity and flexibility 
is afforded by automatic launchers, which fire many 
rockets from each guide. For light rockets like the 
4.5-in. barrage rockets, simple gravity-fed auto- 
matics have displaced multiple launchers in many 
applications. Their disadvantages are (1) the pos- 
sibility of interruption of the salvo by one defective 
round or by improper feeding, (2) a considerable 
increase in dispersion caused by the effect of the 
blast of one rocket on the flight of the following one, 
and (3) a limitation on the quadrant angles at which 
the launchers will operate. Far outweighing these, 
however, are the advantages of decreased weight and 
of standardization; a few miscellaneous fittings en- 
able the same launcher to be used either singly or in 
multiple from virtually any type of vehicle or ship. 
The primary application of multiple-guide launchers 
is for larger rockets or for tactical situations where 
variable train and elevation are required. a 

Finally, launchers may be classified by their type 
of mount, which is determined obviously by the 
tactical use. In certain cases (for example, the CIT 
Type 60 32-barrel closed-breech launcher designed 
for use with 5.0-in. spinners against suicide planes) 
continuous variation in train and elevation may be 
required, and some standard artillery mount has 
usually been used. Such flexibility is usually not 
essential, however, and in the interest of simplicity 
it has been the practice to give launchers as few 
degrees of freedom as possible. Some launchers 
have fixed mounts, set, for example, to fire at 45- 
degree elevation and aimable only by turning the 
vehicle on which they are mounted. Most mounts 
are semifixed, that is, elevation and/or train may be 
adjusted before firing, but not during the firing. 
The required accuracy of adjustment depends on 
the accuracy of the round and the stability of the 
firing platform. 

1722 Blast 

In the design, installation, and use of rocket 
launchers, blast is usually an important problem. 
Although the direct blast is confined to a cone 
narrower than the nozzle exit, it may cover a sizable 

a The Navy Mk 102 launcher is an example of a powered 
automatic, with elevation and train continuously variable 
during firing. 


area at some distance back from the round. Also, 
the air surrounding the direct blast cone acquires 
high velocity. On any obstruction large enough to 
intercept all of it, the blast may exert a force 
roughly equal to the thrust on the rocket — for ex- 
ample, 20,000 lb for the 11.75-in. aircraft rocket. 
The blast can also ignite, burn, or scorch objects 
exposed to it. Hence personnel and equipment, 
including the launcher itself, must be protected 
from blast. The simplest way is to locate the 
launcher where blast need not be deflected, as, for 
example, at the extreme rear or outboard of vehicles 
and boats. When this is impossible, simple blast 
deflectors are used, with small recoil effects. A few 
closed-breech tube launchers have been used, in 
which the gas reverses its direction and escapes for- 
ward around the rocket. In this case the recoil 
forces, though substantial, will not usually rival 
those of an equivalent gun because only a small 
fraction of the propellant burns while the rocket is in 
the launcher. 

In all cases, launcher parts are (or should be) 
designed to expose minimum area to the blast, all 
auxiliary equipment is securely mounted, as far off 
the rocket axis as possible, and electrical assemblies 
are completely enclosed. 

17 2 3 Firing Systems 

Most rockets are fired electrically and require a 
current of at least 3^2 ampere for reliable ignition. 
The components of a firing system are a source of 
power, a control and distribution panel, and the 
sockets or contacts on the guides themselves, to- 
gether with the necessary wiring. Although the 
design problems are mostly straightforward, 3 they 
require careful attention, for failure of the electrical 
system is one of the most common difficulties ex- 
perienced in rocket installations. A storage battery, 
magneto, or blasting machine suffices as a source of 
power. 

Since rockets are almost always fired in salvo, a 
control panel is required. This usually incorporates 
a safety plug, master power switch, indicator lamp, 
push-button firing switch , and individual push but- 
tons or a selector switch for the circuits to the 
launchers. Proper design here is essential to prevent 
accidental firing. The safety plug is removable and 
should be carried by the loader while at the launcher. 
Both it and the firing switch should be double-pole, 


140 


ROCKET LAUNCHERS 


both live and ground leads running through them, 
so that no short circuit or error in wiring can set 
off a round. 

Wiring and contact problems become more com- 
plicated on shipboard because of the possibility of 
deposition of salt from the spray, which may short- 
circuit the contacts if they are designed with im- 
proper clearance, and because of the existence of 
stray potentials, sometimes amounting to several 


relatively few basic elements, only a few aircraft 
launchers have gotten beyond the test stage, and 
there are almost as many basic types as launchers. 
Consequently most of the design problems have 
been specific to a particular type. The launchers 
which have reached service use are described in 
Firing of Rockets from Aircraft. 2 These and one or 
two others are discussed briefly in the remainder of 
this chapter. 



/ 

installation 


under 


port wing of PBY-5, loaded with 


ASR’s. 


Figure 1. Vertical bombing 


volts, arising from galvanic action or from leakage. 
In multiple circuits with parallel wiring, undesired 
ignition may occur through “sneak circuits”; the 
location and elimination of these may be extremely 
time-consuming. The nature of the firing circuit 
depends on whether single shots, ripple fire, simul- 
taneous salvos, or combinations of these are re- 
quired, and the various possibilities are discussed 
in some detail in Rocket Launchers for Surface Use. 1 

17 3 AIRBORNE ROCKET LAUNCHERS 

In contrast to the situation for surface launchers, 
where there is a great multiplicity of designs of a 


17 3 1 Launchers for Retro Firing 

For the attack of submarines from airplanes di- 
rectly above them (as required by the characteristics 
of the magnetic airborne detector [MAD]) CIT de- 
veloped a series of vertical antisubmarine rockets 
[VAR] known also as retro rockets or retro bombs. 
In use, these were mounted under the airplane 
wings (usually) and projected backward with speeds 
just sufficient to cancel the forward speed of the 
airplane, to fall vertically. The launchers adopted 
for service use with these rockets consisted of chan- 
nels about 7 in. wide, 2 in. deep, and 8 ft long, fabri- 
cated of Dural sheet. The rockets, with heads 



AIRBORNE ROCKET LAUNCHERS 


141 


and tails 7.2 in. in diameter, were provided with 
lugs which engaged the lower edges of the channels. 
An electrically insulated spring latch at the forward 
end of the launcher made electrical contact with an 
insulated ring on the tail and held the rocket in 
place until fired. The launching channels, in the 
form of inverted troughs above the rockets, pro- 
vided blast protection for the airplane. 


top. Both 4.5-in. and 7.2-in. rockets were tested. 
The other was a Dural framework for launching 100- 
lb bombs backward from under the belly of the 
A-20C, using six 2.25-in. rocket motors for propul- 
sion. Results of tests of these installations are 
described in the PMC 7 and NMC 8 series of CIT 
weekly progress reports for the period from October 
1942 to May 1943. 



Figure 2. Drawing of Mk 4 launcher, showing the method of mounting and the harmonization adjustment. 
Rocket shown is the 3.5-in. AR Model 1. 


The method of attachment of the launchers was 
necessarily different for each type of airplane. In 
all cases they were mounted in groups of from 4 to 
12 and so oriented as to obtain the desired impact 
pattern. The only two installations which saw 
service use were those on the PBY-5 (see Figure 1), 
which had twelve launchers under each wing, and 
the TBF-1 , which had a total of 8 launchers mount- 
ed on the outside of the bomb bay doors. b 

A considerable amount of experimental work was 
done on two other installations, neither of which 
reached the stage of service use . One was an instal- 
lation consisting of one or more tubes passing 
through the fuselage at an angle so that the faster 
rockets were directed downward and backward and 
the blast came out the open end of the tube at the 

b Reports on the retro installations include references 4, 5, 
and 6. 


17 32 Rail Launchers for Forward 
Firing 

When designs for forward-firing launchers were 
first discussed, it was realized that finners launched 
in a high-velocity head wind would have so little 
dispersion that it might be possible to have the 
launchers extremely short. The British were using 
long slotted rails mounted under the wings for 
forward-firing their 3.0-in. aircraft rockets, how- 
ever, and had reported trial and abandonment of 
very short launchers. It was decided that the rela- 
tively long British-type launchers held the most 
promise for first quick development. The result of 
this decision was the Mk 4 so-called “T-slot” 
launcher shown in Figure 2. It is a Dural box 90 in. 
long with a %-in. wide slot running the full length 
of the lower side to engage the two lugs on the top 


142 


ROCKET LAUNCHERS 


of the 3.5-in. aircraft rocket. A spring catch at the 
back permits breech loading and prevents the rocket 
from sliding out to the rear. Forward of this catch 
is a shear-wire latch which is held by a copper shear 
wire strong enough to retain the round during 
arrested landings but weak enough to shear under 
the thrust of the rocket when fired. Electrical con- 
tact to the round is provided by a two-prong plug 
at the rear of the rail. Later the length of the rail 
was reduced to 70 in. for some airplanes. 

1733 Post Launchers for Forward 
Firing 

The decision to try rail launchers first was an 
unfortunate one, because within a year from their 
first use they were abandoned, leaving the rockets 
with lug bands which, designed for the T-slot rail, 
were far from ideal for the launchers subsequently 
designed. After five months of work on rail launchers 
experiments with “zero-length” launchers began, 
and it became evident quickly that the loss in 
accuracy was only about 2 mils in most cases, not 


other. The short slotted rail of the front one (Figure 
3) engages the standard sway-brace type lug band 
which was originally designed for the long rail 
launcher; a flat tongue on the rear one (Figure 4) 
fits into a tunnel-type lug band which is lower than 
the front lug to ensure clearance of the front post 




Figure 3. Front post of post launcher with Mk 1 
HVAR motor attached. 

enough to justify the greater weight, complexity, 
and drag of the rail launchers. 

The basic design of all zero-length launchers (later 
officially designated “post launchers”) is the same 
except for the latch mechanisms. Each launcher 
consists of two posts under the wing, one behind the 


Figure 4. Rear post of post launcher with Mk 1 

HVAR motor attached and shear wire in place. 

when the rocket is fired. The fuze-arming solenoid 
is usually in the front post and a latch and electric 
receptacle in the rear. Various types of latches for 
holding the round on the launcher were designed, 
and at the end of World War II no decision on the 
most satisfactory type had been reached. The 
rocket becomes free immediately after firing, the 
front lug being guided for only l l /2 in. and the 
rear for about Yi in. 

The first standard production post launcher was 
the Mk 5 , die-formed and spot-welded from aluminum 
sheet. This worked well with the 3.5- and 5.0-in. 
aircraft rockets (50 and 80 lb) but it was insuffi- 
ciently strong for the 140-lb, 5.0-in. high-velocity 
aircraft rocket [HVAR]. SAE 4130 alloy steel 
proved to be the best material and was specified 
for the launchers designed later for Army planes. 

On most airplanes several of these launchers were 
mounted under each wing in front of the landing 
flaps, since it was thought that the usual minor 
blast damage occurring at thie trailing edges of the 


AIRBORNE ROCKET LAUNCHERS 


143 


wings could be withstood better by the flaps than 
by the ailerons, which are weaker and more critical. 
On the P-38, however, the propeller circles cover 
almost the entire flap region, and the only space 
available for mounting launchers was in front of the 
ailerons, where the problem was complicated by the 
fact that the wing chord is too small for normal 
fore and aft separation of vertical post launchers. 
To meet this requirement, the tree-type launcher 
(Figure 5) was developed. This consists of one 
pair of large posts under each wing; each post has 
branches which permit the loading of five rockets in 
a cluster. 

All these launchers will accommodate the 3.5-in. 
AR’s, the 5.0-in. AR’s and the 5.0-in. HVAR/s, 
since these rockets have the same type and spacing 
of suspension lug. For the 2.25-in. subcaliber prac- 
tice ammunition, which have lug buttons only about 
18 in. apart, a special adapter (Figure 6) was de- 
veloped which is mounted on the post launcher in 
the same way as a standard full-caliber rocket, and 
is held in place by a shear wire which is approx- 



Figure 5 A. 5.0-in. HVAR’s loaded on tree 
launcher under wing of P-38L. 



Figure 5B. 5.0-in. HVAR’s loaded on tree launcher under wing of P-38L, 


144 


ROCKET LAUNCHERS 



FRONT SUSPENSION BUTTON 


REAR LAUNCHER STUD 


COPPER SHEAR WIRE 


LAUNCHER LATCH 


STEEL SHEAR WIRE 


Figure 6. Mk 6 launcher attached to post launcher and carrying SCAR. 


imately twice as strong as that used to secure the 
subcaliber round itself.® 


17 34 Launchers for Large Aircraft 
Rockets 

The history of development of launchers for the 
11.75-in. aircraft rocket (“Tiny Tim”) parallels in 
one respect that for the smaller forward-firing 
rockets — the most complicated type was inves- 
tigated first and later abandoned for the simplest 
type. When Tiny Tim appeared on the horizon in 
the late spring of 1944, consideration was given to 
three methods of launching it: (1) displacement 
launchers which swing the rocket away from the 
airplane while holding it parallel to the line of flight 
and fire it automatically when it reaches the 
maximum separation; (2) drop launchers which 

c More complete information on all of these launchers is 
available in many reports listed under Contract OEMsr-418 
in the general bibliography in the appendix. 


release the rocket and ignite it subsequently by 
means of a lanyard or time delay device; and (3) 
fixed launchers which fire the rocket from the 
carrying position in the same way as for smaller 
rockets. It seemed fairly certain that fixed launchers 
of acceptable length and drag would not provide 
enough separation to prevent blast damage; subse- 
quent tests showed this to be the case except for a 
few very rugged aircraft. The drop launcher was 
mechanically simple but was expected to give larger 
dispersions and hence required detailed study of its 
ballistics; thus it promised to be a relatively long- 
term development. The displacing gear appeared 
to offer the best possibility of being put into service 
quickly, since it promised to provide adequate 
separation without loss of accuracy. 

The displacement launcher consists essentially of 
two parallel arms of equal length, with their upper 
ends attached to the two pivots on (or in) the air- 
plane. A latch on the rear arm engages a lug on the 
rocket; to brace against side sway, latches on the 
front arm hold the rocket by two lugs on the for- 


AIRBORNE ROCKET LAUNCHERS 


145 


ward lug band. After the rocket is mounted on the 
launcher, it is swung up and back to the carrying 
position, where it is held by a standard bomb rack. 
In operation, the bomb rack is tripped, and the 
rocket swings down on the launcher arms by 
gravity. As the arms approach the bottom of their 
swing, they actuate a microswitch to ignite the 


ical and structural upkeep problems; (2) inactivation 
of the bomb bay for other purposes when installed 
internally (as on the TBF) or (3) excessive air drag 
when installed externally (as on the F4U); and (4) 
interference with sighting because the centrifugal 
force of the rocket in its circular path causes a 
pitching of the airplane which was as much as 20 



Figure 7. Drop launching of 11.75-in. AR from F6F. 


rocket motor and by cam action release the lug 
latches. The rocket propels itself off the launcher, 
which is stopped in its swing by snubbing cables. 
The airstream throws it back to a horizontal posi- 
tion, where it is retained by a latch. 

After certain mechanical “bugs” were worked out 
of them these launchers were satisfactory on rugged 
aircraft such as the F4U. In general, however, they 
had the following disadvantages: (1) major mechan- 


mils for lightweight fighter craft. Some thought 
was given to the design of a launcher with controlled 
displacement — one that would lower the rocket to a 
safe firing distance and hold it there rigidly while it 
was aimed and fired. None were tested, however, 
because of the success of the drop launcher, the 
next development. 

One squadron of F4LPs was equipped with dis- 
placement launchers and trained in their use, but, 


146 


ROCKET LAUNCHERS 


before it went into combat, the success of the drop 
launcher had made its equipment obsolete, and the 
squadron was recalled. 

In drop launching the aircraft releases the rocket 
as a free body, whereupon it moves under the in- 
fluence of gravitational and aerodynamic forces, 
connected to the airplane only by the nonrestrain- 
ing cords necessary to ignite it. Ignition occurs by 
the closing of a lanyard switch when the cords have 
reeled out the appropriate separation between rocket 
and aircraft. Figure 7 is a series of pictures taken 


from the PBJ equipped with launchers supported 
by a cantilever structure on both sides of the fuse- 
lage just outboard of the bomb bay. The F4U 
installation is shown in Figure 8. 

During the last few months of World War II, 
some experimental work was done by CIT with 
fixed launchers for the SB2C and the P-47. The 
former experienced rather severe buffeting by the 
blast, but five pairs of rounds were air-fired from 
the latter with only slight disturbance to the flight . 
In both cases the centerline of the rocket was spaced 



Figure 8. Two Tiny Tims on pylon drop launchers of F4U-1D. Note practice heads. 


during drop launching of Tim from an F4U. The 
rounds are supported in standard bomb shackles; 
in some cases it is possible to use the bomb release 
equipment supplied with the airplane. Usually, 
however, special sway bracing and other supporting 
structure is required. Before the end of the CIT 
work 11.75-in. rockets had been successfully drop- 
launched from the following airplanes: the F6F, 
F4U, SB2C, F7F, P-38, and P-47 using standard 
bomb stations; from the A-26 using a special struc- 
ture which held two rounds in the bomb bay; and 


approximately 2 ft from the wing. Because such 
a launcher gives smaller gravity drops and sight 
corrections than does the drop launcher, further 
work on it appears to be definitely worth while. 

1 . 7 . 3. 5 Aircraft Launcher Design 

Problems 

In all aircraft installations the same basic prob- 
lems naturally arise: attachment of launcher to air- 
plane, attachment of round to launcher, firing cir- 


AIRBORNE ROCKET LAUNCHERS 


147 


cuit design, and blast problems. Their solution 
varied so much on different aircraft that generaliza- 
tion is difficult, and the weekly progress reports 
must be consulted for details. These problems are 
discussed in considerable detail in Firing of Rockets 
from Aircraft ; 2 we shall mention only that of blast. 

Damage to aircraft from rocket firing results from 
three causes: (1) ejected material from the rocket 
motor (closure disks, igniter wires, drying bags, 
etc.), (2) a shock wave from the firing of the igniter, 
and (3) the turbulent high-velocity airflow induced 
by the rocket jet. In no case was a rocket placed 
so that the jet itself impinged on any part of the 
plane (except for a few very small rockets fired 
from closed-breech tubes). For small rockets (i.e., 
5.0-in. and smaller) the leading edges of the wings 
and stabilizers suffer the most damage from ejected 
material, and the trailing edges of flaps and ailerons 
are the parts most subject to damage from the com- 
bination of causes (2) and (3) . Exposed fabric sur- 
faces near the blast usually require light metal 
sheathing, and internal reinforcement is sometimes 
found necessary. 

Tiny Tim, with its 150 lb of propellant and its 
luminous jet over 100 ft long, naturally gave blast 


problems of much greater severity. Anyone who is 
close to one when it is fired is likely to acquire a 
permanent feeling of amazement that such a rocket 
could be launched from aircraft at all. After one 
plane crashed immediately after firing, many weeks 
were consumed in tests to investigate the blast 
effects on various aircraft. It was shown eventually 
by elaborate high-speed photographic tests that the 
igniter shock wave was doing most of the damage, 
and a reduction of the igniter charge to the bare 
minimum consistent with good ignition removed 
most of the difficulty. These tests are discussed in 
detail in reference 9. 

17 ' 3 ' 6 Launchers for Aircraft Spinners 

The development of aircraft spinners was still in 
its infancy at the termination of the CIT work, and 
little attention had been paid to launcher designs. 
The launchers that were used in tests were essen- 
tially identical with the ground launchers, often 
attached with lug bands to the regular post launch- 
ers for finners, and probably have little similarity 
to the launchers which will be designed for service 
use to exploit the peculiar advantages of spinners. 


Chapter 18 

SERVICE DESIGNS OF FIN-STABILIZED ROCKETS 
FOR SURFACE WARFARE 

By C . W . Snyder 


i8i INTRODUCTION 

I n chapters 18, 19, and 20, we shall discuss 
briefly each of the rockets which were developed 
by Project OEMsr-418 and which either were used 
by the Services in World War II or which had a 
significant influence on the design of later rockets 
which were used. In each case, we shall indicate the 
service requirements which the rocket was intended 
to meet and sketch the reasons which impelled the 
choice of particular designs to meet them. In some 
cases we may be able to evaluate the success of the 
rocket in combat, but relatively little information 
on this point is available find Navy or Army files 
must be consulted. 

It is intended that Chapters 18, 19, and 20 be 
read in connection with the following three volumes 
published by CIT as part of the final report of the 
project: Ballistic Data, Fin-Stabilized and Spin- 
Stabilized Rockets , l which contains photographs, 
weights and dimensions, and interior and exterior 
ballistics data for virtually all rockets mentioned in 
these chapters; Rocket Launchers for Surface Use, 2 
which contains photographs, description, and bibli- 
ography on every surface launcher which was used 
outside of the project itself and which the project 
had a hand in developing; and Firing of Rockets from 
Aircraft: Launchers, Sights, Flight Tests, 3 which, in 
addition to much other information, includes in the 
first chapter short descriptions and photographs of 
all service airborne launchers which CIT aided in 
developing. 

182 ANTISUBMARINE ROCKETS [ASR] 

The antisubmarine rocket [ASR] was the first 
American rocket to “go to war.” Tests on a similar 
projectile began at CIT in January 1942. The 
actual birth of the ASR, however, was in a meeting 
of March 7, 1942, between representatives of Divi- 
sions C and A, NDRC. There it was decided that 
the projectile should be similar to the British 


“Spigot Gun” or “Hedgehog,” except that it was 
to carry 40 lb of TNT in a total weight of 80 lb for a 
range of 200 yd. The rockets were to be fired in 
salvos of four or six so as to have a separation of 
about 20 ft on striking the water. There was sore 
need for such a weapon because investigation of 
records in Germany following World War I had dis- 
closed that the conventional type of depth charge 
attack had been not nearly so effective as had been 
assumed, the principal reason being that sound 
contact with the submarine cannot be maintained 
at close range, and, during the interval after contact 
is lost but before the ship is close enough to begin 
its attack, effective evasive action can be taken by 
the submarine. Also, after the depth charges have 
exploded, the water is so full of echoes that it is 
seldom possible to regain sound contact. 

Thus the requirements were as follows: 

1. Range great enough so that the submarine 
could be attacked while maintaining sound contact 
with it. 

2. Dispersion small enough so that a predeter- 
mined shot pattern could be laid down, calculated to 
give the highest probability of a hit. 

3. Projectile to be capable of launching from 
small boats. 

4. Payload great enough so that a single direct 
hit could inflict lethal damage on a submarine . 

5 . Contact fuzing so that sound contact need not 
be severed by an explosion unless a direct hit is 
scored. 

The British had developed for this purpose a for- 
ward-thrown projectile called the Hedgehog from 
the fancied resemblance of its launcher to the ani- 
mal with its spines bristled up. The launcher con- 
sisted of a group of steel rods inclined at forward 
angles and welded to the deck of a ship. The pro- 
jectile itself looked almost exactly like the final ASR 
on the exterior, and its propelling tube, carrying 
stabilizing fins, slipped over the steel rod. Propul- 
sion was provided by a charge of black powder in 
the forward end of the tube. It was desired to 
improve on the Hedgehog in three respects: (1) by 


148 


ANTISUBMARINE ROCKETS [ASR] 


149 


eliminating recoil so that the projectile would be 
usable on small boats, (2) by increasing the payload, 
and (3) by increasing the range. 

Twenty-three days after the meeting which au- 
thorized the rocket, firings from shipboard at sea 
were made. Everything about the test was satis- 
factory except that it was found that 85-lb projec- 
tiles were difficult to handle on the rolling, pitching, 
spray-drenched deck of a small ship. It was there- 
fore decided to copy the Hedgehog almost exactly 
in weight and shape. Such a rocket was standardized 
almost immediately, and no further significant 
changes were made on it except the substitution of a 
tubular three-ridge charge for the original tubular 
charge with celluloid spacers. Its nose-fuzed head 
is 7.2 in. in diameter and 19 in. long and has only 
a 0.10-in. wall thickness, so that 30 of its 50-lb un- 
fuzed weight is TNT. A later head, the Mk 5, has a 
still thinner wall. The 2.25-in. 11-gauge motor is 
equipped with a machined screw-in nozzle and a 
ring tail 7.0 in. in diameter. Photographs and 
drawings of the rocket are given in Ballistic Datal 

182 1 Designation and Types 

In the early days when “rocket” was a restricted 
word, the ASR was known as the antisubmarine 
bomb [ASB], and for a time the official Navy desig- 
nation was Antisubmarine Projector Mk 20, Charge 
for [ASPC]. More often than not, however, the 
rocket was referred to even in official communica- 
tions as the “Mousetrap,” the name deriving from 
the appearance of the launcher, which folded flat 
against the deck when not loaded. These designa- 
tions applied loosely to any of the rockets in the 
ASR series. 

In order to give the same velocity (175 fps) to two 
different heads (the one originally designed by CIT 
and the British Hedgehog head which was finally 
adopted as standard in the interest of uniformity) 
and three different fuzes (the HIR 1 or Mk 135, 
the HIR 3 or Mk 140, and the British-designed 
underwater-vane-arming fuze Mk 131), four dif- 
ferent grains were designed, all having an outer 
diameter of 1.70 in. and a length of approximately 
1 1 .6 in. , but differing in perforation diameter to give 
weights of 1.40, 1.43, 1.50, and 1.55 lb. Nine dif- 
ferent combinations were originally distinguished 
by complete round Mark numbers, but in the lat- 
est revision of nomenclature, all ASR’s are desig- 


nated 7.2-in. Rocket Mk 1 Mod 0. The various 
combinations of components are given in Ballistic 
Data. 1 


Launchers 

The first launcher used, and the one which gave 
the Mousetrap its nickname, was the 7.2-in. Type 4 
launcher, designated Mk 20 by the Navy. It had 
four formed steel rails spread slightly apart to pro- 
vide a suitable shot pattern. It could be folded flat 
for stowage but was not adjustable in quadrant 
angle since maximum range was desired. A pho- 
tograph of it is included as Figure 1 . Later the Navy 



Figure 1. Mk 20 “Mousetrap” launcher loaded 
with ASR’s. 

designed a similar double-deck launcher, the Mk 22, 
for eight rounds. Both were used extensively, 
mounted in pairs on the foredecks of PC boats, 
Coast Guard cutters, harbor patrol vessels, de- 
stroyer escorts, and other types of vessel. 

18 2 3 Design Features 

Nozzle. The ASR was designed hurriedly and 
before any extensive investigation of nozzle types 
had been made, and a machined threaded nozzle 
was chosen because it could be made easily and 


150 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


accurately. It is a good nozzle but is somewhat 
more costly than the formed nozzles that were sub- 
sequently developed. 

Tail. A ring tail was chosen simply because the 
British had used it, but underwater tests later 
showed it to be a good choice. Much work on 
underwater ballistics of the round was done by the 
Morris Dam group at CIT (see Part I of this vol- 
ume) , and as a result of their findings two changes 
were made in the tail. 4,5 The four radial vanes sup- 
porting the tail rings were canted at a 10-degree 
angle, imparting a slow spin to the rocket and re- 
ducing “ wandering” under water, and the rings 
themselves were streamlined, the front edge being 
rounded and the rear end tapered to a sharp edge, 
thus reducing the underwater drag and increasing 
the terminal velocity. The tail ring diameter was 
made less than the head diameter to reduce tip-off a 
which would have been significant on such a slpw 
projectile. Since the center of mass of the whole 
rocket is in the head, the tail does not ride on the 
launcher at all. 

Contacts. The system first used on the ASR 
motor of making electrical contact to the igniter 
was subsequently used on most fin-stabilized rockets . 
The tail shroud is composed of two rings, the rear 
one being welded to the radial fins and the front 
one being insulated. The igniter leads are provided 
with lugs and screwed to small metal angles inside 
the two rings. Spring-loaded knife contacts on the 
launcher make electrical connection with the rings. 

General Shape. A number of alternative shapes 
were tested for underwater behavior by the Morris 
Dam group. Hemispherical noses and noses flatter 
in varying degrees, several tail shapes, streamlining 
the rear of the head, putting an air space in the 
rear of the head to increase the righting moment 
under water — all these were tested. Several of these 
designs gave considerably higher terminal veloc- 
ities than the standard, which would be a decided 
advantage, but none showed any marked improve- 
ment in underwater dispersion and virtually all 
gave greater forward travel after impact than the 
standard. 

Igniters. The original ASR had a brass case 
igniter b with a bakelite disk closure, and a formed 
celluloid “saddle” was cemented to the front end 

a Tip-off is the reduction of the effective launching angle by 
gravity drop of the head while the tail is still constrained by 
the launcher (see Section 24.4.3). 

b See Chapter 22, Figure 13, A and B. Igniters are dis- 
cussed in Section 22.11. 


of the powder grain to hold it in place and prevent it 
from being squeezed between the grain and the 
front closure disk. In a few months the bakelite 
disk was superseded by a molded cellulose acetate 
closure, which provided a much better seal and 
did away with the saddle . Later the molded plastic 
case igniters with screw closures were specified for 
this rocket, as for most others. 

Grains. The original grain was tubular, 11.6 in. 
long, 1.7 in. OD, and 0.6 in. ID. It was spaced in 
the tube by cellulose nitrate strips cemented to the 
grain with Duco household cement. It was found 
that strips gave 0.7 times as much impulse per 
pound as did the ballistite itself, and hence this 
fraction of their weight (and half the igniter weight) 
was included in the “effective weight” of the grain. 
Most of the experimental static-firing tests which 
led to the discovery of the stabilizing effect of 
radial holes c in the grain were made with ASR 
motors, and, as soon as the effect had been proved, 
radial holes were specified for all grains. The idea 
of extruding ballistite ridges on the grain to eli- 
minate the necessity for celluloid strips was also 
tested first on the ASR and then became standard 
practice. Originally a cellulose acetate washer was 
cemented to the grid end of the grain, but it was 
later found to be unnecessary and abandoned. 

The cast iron stool grid (Chapter 22, figure 15 A) 
was originally specified and remained standard 
because the shape of the machined nozzle (Chapter 
23, figure 3 A) does not give sufficient port area 
with the box grid. 

18 2 4 Reports on the ASR 

The very early history of the rocket and its 
launcher is contained in reference 6. Further de- 
velopment is reported in detail in reference 7. In 
particular, the second volume of this report dis- 
cusses the experimental tests which first demon- 
strated that the burning of a tubular ballistite 
grain could be stabilized by the use of radial holes. 
Instructions for use of the weapon in service are 
given in reference 8, and amplified and revised in 
reference 9. A comprehensive study of the factors 
determining the success of antisubmarine attacks 
by Hedgehog and Mousetrap projectiles is given in 
reference 10. See also the reports of the Morris 
Dam group. 11-20 Design of the grain is described in 

“See Section 22.6 and reference 7. 


BARRAGE ROCKETS [BR] 


151 


reference 21 . Reference 22 describes the two service 
launchers. 

18 2 5 Related Rockets 

As the first rocket standardized for service use, 
the ASR naturally inaugurated many design 
features which are found in later rockets. Thus the 
BR, the VAR series, and the SCAR (all of which 
are discussed in detail in the following pages) con- 
tain elements borrowed directly from the ASR. 
The rocket was taken over almost intact for the 

7.2- in. demolition rocket Model 17 [DR] which was 
designed for the Army Engineers. It has the ASR 
motor and a head which is almost identical with 
those of the VAR series but contains the PIR base 
fuze Mk 146. An adapter connects the 3.25-in. 
head threads to the 2.25-in. motor threads. The 
head is filled with plastic C-2 explosive, and the 
rocket is intended for demolition of concrete walls 
and similar obstructions. Ordinarily it is used for 
virtually point-blank fire. Its service designations 
are 7.2-in. Rocket Mk 1 Mod 2 and Rocket, HE, 

7.2- in., T37. It is described in reference 23. 

A service launcher (T-40) was designed by Army 
Ordnance, although CIT assisted in its develop- 
ment. It consists of 20 tubes in an armored housing 
mounted on the turret of an M4A1 medium tank 
and attached to the gun so that it may be aimed 
by using the gun mechanism. Its predecessor, the 
CIT 7.2-in. Type 5 launcher, described in Rocket 
Launchers for Surface Use, 2 was superseded by the 
turret-mounted version because it lacked an inde- 
pendent train adjustment and interfered with the 
tank’s maneuverability on rough terrain. 

The DR is understood to have been used in the 
Normandy landings and the subsequent European 
campaign, but little is known about it at CIT. 

Like the earlier BR, the DR was redesigned to 
give better accuracy by lengthening the motor and 
substituting a thinner-web grain. The fast-burning 
Model 18 has a lateral dispersion of less than 5 mils 
at all temperatures above 10 F when fired at 32 
degrees QE from the 7.5-ft T— 40 launcher. It did 
not get into production for service use and has no 
service designations. 

A short-range DR was also designed for the 
purpose of countermining Japanese J-13 antiboat 
mines by firing ahead of a landing boat . This Model 
19 rocket uses a standard-length motor and a 5-in. 


length of the thin-web grain. Fired at 45 degrees 
QE from the 7.5-ft launcher, its range varies with 
temperature from 80 to 120 ft, and its dispersion is 
2 mils or less. It did not reach service use. 


183 BARRAGE ROCKETS [BR] 

The 4.5-in. barrage rocket [BR], originally called 
beach barrage rocket [BBR], was first suggested on 
June 16, 1942, just a few weeks after the stand- 
ardization of the ASR, and its development pro- 
ceeded rapidly. The first models were test-fired on 
June 24, the first full-scale sea test was on July 28, 
and the first service use was in the assault on Casa- 
blanca on November 8. 

The requirements for the rocket were simple. 
No weapon existed which could fill in the gap of a 
few minutes between the time when the naval and 
air barrage had to be lifted and the time when the 
first invading troops hit the beachhead. This short 
respite from bombardment was enough to allow the 
enemy to organize and pour a devastating fire into 
the landing waves, and casualties in the first wave 
were alarmingly high, as everyone will remember. 
The rocket was intended to be carried on the troop- 
carrying boats themselves and to continue bom- 
barding the beachhead up to a few seconds before 
the actual landing. A light-case head for maximum 
fragmentation and antipersonnel effect and a range 
of approximately 1,000 yd were suggested. Dis- 
persion was of little importance, and in fact a rela- 
tively high dispersion might be preferable, since a 
large area could then be covered without the com- 
plication of having to “fan out” the launching rails. 

18 3 1 Designation and Types 

The original BR design incorporated a pressure- 
arming base fuze, but this was quickly abandoned 
in favor of a point-detonating fuze which detonates 
the head completely above ground and is thus more 
effective against personnel. To accommodate the 
base fuze, the original motors had internal threads, 
and, with its abandonment, the motor was simply 
shortened slightly leaving the internal threads so 
as not to have to change the head design. This 
was the 4.5-in. Rocket Mk 1 Mod 0, consisting of 
the 2.25-in. Mk 7 Mod 0 motor, the 4.5-in. Mk 1 
Mod 0 head, and the original AIR fuze which was 


152 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


a modified PDF M-52 trench mortar fuze. Shortly 
before the end of 1942, the design of both motor 
and head was changed to use external threads on 
the motor, and at approximately the same time an 
improved AIR fuze, the Mk 137, was introduced. 


than any other American rocket. In the latest 
nomenclature, all the above rockets are designated 

4.5-in. Rocket Mk 1 Mod 0. The Mk 145 fuze is 
sometimes used in place of the Mk 137 to provide a 
short delay in firing; the rocket designation is then 



Figure 2. Mk 1 Mod 0 “crate” launcher being loaded with inert 4.5-in. BR. 


The new rockets had the following designations: 
CIT production: 

4.5-in. Rocket Mk 2 Mod 0; 

2.25-in. Rocket Motor Mk 8 Mod 0; 

4.5-in. Rocket Head Mk 2 Mod 0. 

BuOrd production: 

4.5-in. Rocket Mk 3 Mod 0; 

2.25-in. Rocket Motor Mk 9 Mod 0; 

4.5-in. Rocket Head Mk 3 Mod 0. 

The various early design changes can be followed 
in the photographs and drawings of references 24, 
25, and 26. Production of the Mk 3 rocket ran 
into the millions, and it probably saw more use 


Mk 1 Mod 1. A smoke head, Mk 5 or Mk 7, was 
designed much later, and the latest designation for 
the rocket with this head and the Mk 137 fuze is 
Mk 4 Mod 0. 


18 3 2 Design Features 

Motor. The nozzle, grid, grain (the 1 .43-lb Mk 1), 
and motor closures for the first BR were the same as 
those of the then standard ASR, and none were ever 
changed, although changes in all of them were dis- 
cussed at one time or another. The igniter went 



BARRAGE ROCKETS [BR] 


153 


through the same evolution as that of the ASR as 
the state of the art improved. 

Internal threads on the front end of the motor 
tube were first specified because they made the base 
fuze design less complicated, but, when an inves- 
tigation was begun to see whether the rocket’s 
dispersion could be decreased, evidence appeared 
indicating that external threads gave better disper- 
sion, presumably because the motor pressure ex- 
panded the internal threads slightly and loosened 
them, increasing the malalignment. It was difficult 
to be certain about this point, however, and the 
increase in the diameter of the filling hole in the 
base of the head which accompanied the change 
from internal to external motor threads was prob- 
ably a more cogent reason for the new design. 

Heads. The fragmentation heads were originally 
made by hot pressing from standard 4.5-in. pipe, 
and, except for the changes in shape to accommodate 
changes in fuzing and in motors, they remained 
essentially the same. Heat treating to improve 
their fragmentation was soon specified. At one 
time there was a discussion of grooving the heads 
like a hand grenade, but tests showed that the 
fragmentation was not improved thereby. Several 
fragmentation tests were made, 27-29 but nothing 
startling was disclosed and no design changes re- 
sulted. The design of the smoke head was straight- 
forward and involved no special problems. Several 
other special purpose heads were suggested and 
tested but never adopted. 

Tails. The very first BR, which had a PIR fuze, 
also had a combination radial-fin tail and ring tail. 
The fins extended to the corners of the 4.5-in. square 
and were insulated from the ring, which was made 
in four quarters. Thus the ring formed one elec- 
trical contact, and the fins or the body of the rocket 
itself formed the other. This required a larger num- 
ber of pieces than the ordinary ring tail, did not 
apparently decrease the dispersion noticeably, and 
somewhat complicated the launcher problem by 
requiring that the rocket be oriented in a certain 
way. It was quickly abandoned in favor of the 
two-ring design like the ASR, and no further 
changes were made on it except the simplification of 
making two adjacent radial fins from a single piece 
of metal. This became the standard design and was 
used on all subsequent ring tail rockets . 

One other type of tail was thoroughly tested and 
is discussed in a report. 30 It had a single shroud 
ring and the insulated contact was a very short ring 


(about % in.) inside the shroud ring at the rear. 
This tail was extremely simple in design and worked 
well, the objection to it being that it somewhat 
complicated the launcher contact problem. Had it 
been adopted, the later development of the auto- 
matic launcher would have been seriously hampered. 

183 3 Accuracy 

Although the 25- to 85-mil dispersion of the BR 
was adequate for its primary purpose of beach 
barrage, there was continual pressure to improve 
it so that the rocket would be better suited to other 
uses. For this reason, and also because the BR 
was a convenient test rocket for learning more 
about the general problem of dispersion (since it 
was inexpensive to make, its heads, when plaster- 
filled, were reusable almost indefinitely, and its 
dispersion was relatively sensitive to changes in 
motor design), a comprehensive program to im- 
prove its dispersion was undertaken and continued 
for several months. 

The first attempt was to find a nozzle and grid 
combination which would give lower dispersion than 
the machined screw-in nozzle and the cast three- 
legged grid and which, incidentally, might be easier 
to fabricate. A considerable variety of nozzle 
shapes were tried along with various methods of 
holding the nozzles in the tube. No combination 
was found which gave significantly less dispersion 
than the standard, and some gave surprisingly large 
dispersions. In all cases the mechanical malalign- 
ments were known and could be corrected for, and 
in several tests the malalignments were made so 
small that they can be ignored. Despite the care 
with which the experiments were done, it is perhaps 
possible that, if they had been repeated two years 
later after better techniques of making formed 
nozzles had been developed for the aircraft rockets, 
the formed nozzles might have shown up more 
favorably in comparison with the machined. The 
various kinds of nozzles tried are described and the 
results are analyzed in references 31, 32, and 33. 

Another line of attack was to try to reduce the 
gas malalignment 41 by straightening out the gas 
flow, running it through long tubes and screens and 
baffles of various types. None of them improved 
dispersion, and some made it much worse. Some 

d For definitions of mechanical malalignment and gas mal- 
alignment see Sections 21.4.1 and 24.8. 


154 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


experiments on spinning the BR, still using fin 
stabilization, were made, but the mallaunching 
apparently nullified the improvement in dispersion 
that might have been obtained. All the various 
expedients tried during this investigation are de- 
scribed in reference 34. 

The final conclusion that the BR gas malalign- 
ment could not be reduced left open only one way 
to increase the accuracy — by decreasing the burn- 
ing time. This meant either operating the motor at 
a considerably increased pressure or using a thinner- 
web grain. The use of high-strength heat-treated 
steel would have permitted higher operating pres- 
sures, but this was felt not to be desirable for pro- 
duction reasons. A safety valve for the front end of 
the motor was tested, which would allow the use of 
higher pressures over the normal operating tem- 
perature range by opening and reducing the pres- 
sure at high temperatures. As long as the valve 
remained closed, the dispersion was actually de- 
creased by the smaller burning time, as expected. 
When the valve opened, however, the gas escaping 
from it had a large “gas malalignment/ ’ and the 
dispersion was poor. The complexity of the valve 
was another argument against it, and its use was 
never recommended. The thin-web grain was suc- 
cessful, however, and was recommended for service 
use. 


Launchers and Service Use 

The heading of this section could well serve as 
the title of a rather large book. Rocket Launchers for 
Surface Use 2 discusses eleven different launcher 
types which were designed by CIT for the 4.5-in. 
barrage rocket and by no means exhausts the list. 
A few additional launchers were designed by the 
Bureau of Ordnance, and a few rockets are known 
to have been fired in combat from makeshift wooden 
launchers nailed together on the spot. At least 
eight authorized launchers saw some service use, 
the most important of which were the 12-rail 
“crate” launcher and the “automatic.” 

The “crate,” CIT 4.5-in. Type 5 launcher, was 
designated Mk 1 by the Navy, Mod 0 being for the 
port side and Mod 1 for the starboard. As shown in 
Figure 2, it consisted of twelve rails connected to- 
gether into a boxlike structure and mounted on 
trunnions to allow adjustment of quadrant eleva- 
tion from 0 to 45 degrees. The rails were 5 ft long 


(plus 1 in. for electrical contacts) and were con- 
sidered the “standard” BR rails for purposes of 
measuring range, dispersion, etc. 

The first service use of the crates was in the 
invasion of North Africa in November 1942, and 
thereafter they were regularly used in landing 
operations in the Mediterranean and European 
Theaters of Operations. They were introduced in 
the Pacific by the Second Engineer Special Brigade, 
which for several months had the distinction of 
being the only rocket unit in that part of the world . 



Figure 3. Mk 7 “automatic” launcher loaded with 
various types of BR: Top to bottom: Mk 4 Mod 0 
rocket (smoke), fast-burning motor with standard 
head, standard motor with incendiary head, four 
standard Mk 1 Mod 0 rockets. 

They first used the weapon in the fighting around 
Finschhafen on New Guinea in October 1943, and 
they spearheaded their first amphibious landing at 
Arawe two months later. For the next six months, 
this group with the crates and four rocket DUKW’s 
had a part in nearly every important landing 
operation. The complete story of their operations 
is a very long one, and it has been told by the 
commanding officer of the 2nd ESB, Brigadier 
General William F. Heavey, USA, in two articles. 35 
A popularized account of the activities through 1944 
is given in The Yale Review.™ The crates were also 


BARRAGE ROCKETS [BR] 


155 


extensively used on PT boats until automatic 
launchers, and later 5.0-in. spinners, became avail- 
able, and they accounted for a large amount of 
enemy shipping. 

The first automatic launcher (Figure 3) was made 
at CIT in April 1943, and its important advantages 
were immediately recognized: 

1. Light weight. For twelve rounds, its weight 
is only 115 lb as compared to 350 for the crate. 


primary advantage. In Rocket Launchers for Surface 
Use , 2 3 4 seventeen typical installations are listed, and 
how many more actually existed is probably im- 
possible to determine. With this launcher, an 
LCM— 3, for example, could fire a ripple salvo of 576 
rockets — a total of 12,000 lb of payload laid on the 
target in about 4 seconds if desired. Even the lowly 
jeep carried heavy artillery as shown in Figure 4. 

The automatic, designated Type 8 by CIT, be- 



Figure 4. Mk 7 launcher installation on jeep. 


2. Simplicity. On electrical wiring, for example, 
the automatic requires only one-twelfth as much 
as other launchers, since a single set of contacts 
fires the whole salvo. 

3. Greater safety for the operator, since it can 
be loaded from the side, whereas most other mul- 
tiple launchers require either breech or muzzle 
loading. 

4 . Adaptability to a great variety of installations. 

Its almost universal adaptability is the launcher’s 


came the Navy Mk 7 launcher, and production by 
the Services ran into tens of thousands. It is 
probably no exaggeration to say that the 4.5-in. 
barrage rocket was the most important rocket used 
in World War II and the Mk 7 was the most im- 
portant launcher. It was this combination that 
helped to teach the Japanese that they could not 
defend a beach and resulted in virtually no opposi- 
tion being offered to the initial wave in the landings 
during the last year and a half of World War II. 


156 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


1835 Reports 

There are a large number of reports on the pro- 
cedure for using the BR with various launchers, 
and for these the reader is referred to the bibliog- 
raphies accompanying the descriptions of the launch- 
ers in Rocket Launchers for Surface Use. 2 References 
24, 25, and 26, although also intended as service 
manuals, include sufficiently complete descriptions 
of the rocket as it was at the time of their writing 
to give a picture of the various steps in its develop- 
ment. Photographs of the very earliest round and 
the original crate launcher may be found in refer- 
ence 37. Manufacturing methods used in CIT 
pilot production are described in reference 38. Tests 
made in developing the fast-burning grain are dis- 
cussed in reference 39. Some tests carried out to 
learn about the suitability of the BR for para- 
troopers’ use are described in references 40 and 41. 


that it was not thought desirable to introduce the 
change. The principal production difficulty would 
have been to change the fuze, which was the critical 
item. The much shorter burning time of the new 
model required extensive modifications of the AIR 
fuze. 

A 250-yd barrage rocket was also developed for 
possible use in detonating land mines. The thin-web 
charge was used to keep the burning time short, 
since a dispersion of 5 mils or less was desired. 
Standard length motors were used even though the 
grain was less than half the standard length, be- 
cause tests showed that the rocket had insufficient 
stability with a shorter motor and gave bad dis- 
persion. The combination of short burning time 
and low velocity gave a relatively low dispersion, 
less than 8 mils at medium and high temperatures, 
but still considerably above the desired value. The 
request for this rocket was withdrawn before devel- 
opment work was entirely complete. 


18 3 6 Related Rockets 

The CIT answer to requests for better accuracy 
with the BR was the so-called “fast-burning” or 
“short-burning” BR. Very little development work 
was required on this round. The perforation in the 
grain was simply enlarged to give a web thickness 
of 0.4 in. instead of 0.55 in., and grain and motor 
tube were lengthened to bring the propellant weight 
back up to 1 .43 lb. The long thin grain gave a rather 
severe drop in effective gas velocity e at high tem- 
peratures, but the reduction in range was only 20 
yd at 115 F, and the upper temperature limit was 
sufficiently high. The comparison of this rocket 
and the standard with regard to lateral dispersion 
at 45 degrees QE is shown in Table 1 . 

Table 1 

Burning time Lateral dispersion 

Temperature (seconds) (mils) 

(degrees) Standard Fast-burning Standard Fast-burning 


10 

0.66 

0.38 

85 

48 

70 

0.37 

0.22 

45 

20 

120 

0.23 

0.14 

25 

4 


This rocket was recommended for service use but 
was never adopted because, by the time it was 
ready, production on the standard model was well 
under wa y and the rockets were needed so urgently 
® See Section 21.1.1. 


184 CHEMICAL WARFARE 

ROCKETS [CWR] 

Development of a rocket for the Army Chemical 
Warfare Service was one of the earliest projects 
tackled by the CIT group, the first field firing 
being on December 23, 1941. The intention was to 
develop a rocket to replace or supplement the 
Livens projector bomb, since the lack of recoil 
would permit the launcher to be mounted on a 
truck, eliminating both the weight of the Livens 
mortar and the time required to emplace it. The 
original specifications called for a projectile to 
carry a liquid payload of between 20 and 30 lb with 
a maximum range in excess of 3,000 yd. No definite 
specifications as to dispersion were made, but it was 
indicated that a dispersion of the same order as 
that of the Livens (probable error 50 yd in range 
and 25 yd in deflection) would be acceptable. Fol- 
lowing the first tests of the projectile at Edgewood 
Arsenal, Maryland, more definite specifications were 
outlined, calling for a bomb of 2.2-gal capacity to 
carry 20 lb of chemical agent for a maximum range 
of 3,400 yd or more. 

The first rocket designed had a motor which was 
patterned closely after the British 3.25-in. motor, 
in that it had a formed nozzle of almost identical 
shape, sealed at the front end by an obturator cup 
and held in the motor tube by a piston ring at the 


CHEMICAL WARFARE ROCKETS [CWR] 


157 


rear (the British RP— 3 uses rivets), a six-legged 
stool grid seated on a nozzle ring, and four radial 
fins. The motor was 3 ft long and extended clear 
through the center of the 7.0-in. diameter head to 
make a compact, streamlined-looking projectile. 
Because of its short length, the rocket had a disper- 
sion of more than 100 mils and was unacceptable 
also because the re-entrant design precluded the 
use of a burster tube in the head to disperse the 
contents upon impact. Moving the motor back so 


Two changes in design were then made: sub- 
stituting a box grid for the complicated cast stool 
grid, an improvement which became permanent, 
and forming the nozzle in one piece with the tube 
instead of inserting it. It was found that the one- 
piece nozzles deflected considerably under hydro- 
static pressure, 42 but no clear evidence of increased 
dispersion from this source was discovered. 5 

On May 29, 1943, a meeting of the JNW Rocket 
Board decided, in the interest of uniformity of 



Figure 5. Launcher, Rocket, Multiple Artillery, 7.2-in., T32, mounted in truck and loaded with 24 CWE-N’s. 


that it was completely behind the head reduced the 
dispersion to about 40 mils and necessitated adding 
a 7.0-in. ring to the tail to support it on the launcher. 
The ring was made in two parts to provide electrical 
contact as with the ASR and BR, but the radial 
fins were left unchanged, extending forward and 
radially beyond the ring. Tests with inertia-firing 
fuzes (both nose and base) indicated that a point- 
detonating fuze was required to give sufficiently 
rapid dispersal of the load. Except for the lack of a 
suitable fuze, the rocket was regarded as satis- 
factory for the contemplated service use. 


design and interchangeability of auxiliary equip- 
ment, to increase the head diameter from 7.0 to 7.2 
in. and to use the Mk 3 VAR (retro rocket) motor 
(see Chapter 19) with a different grain for the CWR. 
The new model was dubbed the CWR-N (N for 
“new”), and it showed a lateral dispersion of 
approximately 66 mils as compared to 40 for the old 
model. Radial fins were therefore added to the ring 
tail, and the dispersion returned to its former value 
(see Figure 11B of Chapter 23). 

{ See reference 43 for comparison of dispersion of various 
models of CWR. 



158 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


18,4 1 Design Features 

Grain. As was the case with all early rocket 
motors, much trouble with the grain was ex- 
perienced in the beginning, and a number of things 
were tried until static-firing tests on the ASR solved 
the problem. A 2.5 x 1.0-in. tubular grain with 24 
radial holes was then adopted, and, except for the 
addition of three ridges, this remained the standard. 

Fuze. The Mk 137 BR fuze was used on the CWR 
for a time, but the large propeller was not necessary 
for such a fast projectile and the large protective 
cup surrounding it reduced the range by 100 yd. 
The Mk 147 fuze with a much smaller propeller was 
developed especially for the CWR. In accordance 
with standard practice for chemical bombs, the 
fuze detonates a burster tube which extends virtu- 
ally the full length of the head. 

Nozzles and Accuracy. Although satisfactory in 
other respects, the CWR suffered from the usual 
ailment of rockets — insufficient accuracy. As pre- 
viously mentioned, the accuracy was improved on 
two occasions by tail changes which increased the 
stability. At several times during the long period of 
development of the CWR, tests of various nozzle 
changes were made in an attempt to decrease dis- 
persion. The first nozzle had an abrupt entrance 
cone similar in contour to that of the machined 
ASR nozzle (see Figure 3 of Chapter 23), and this 
was changed to a more gradually tapering entrance 
on the basis of yaw machine tests which showed 
that longer nozzles gave smaller side forces. It is 
probably impossible to draw any conclusions from 
the data on the effect of the later change from insert 
to integral nozzles, because the observed dispersions 
varied so widely from test to test and various other 
factors were being changed from time to time. It 
was thought that the reduction in nozzle expansion 
ratio entailed in the change to the YAR-type motor 
might increase dispersion, but the field tests of this 
point gave negative results. 44 The difference in 
dispersion between rockets having nozzles with 
smooth and rough interior finishes was found to be 
so small that it could not be clearly separated from 
the malalignment effect. It was thought that the 
orientation of the nozzle throat, or more accurately 
of the portions of the entrance and exit cones close 
to the throat, might be more important than the 
orientation of the exit cone which was normally 
assumed to define the direction of gas flow. Tests 
indicated that “throat malalignment’’ does have an 


influence on dispersion but that it is less important 
than the ordinary “mechanical malalignment.” 45 
The only change in nozzle design which ever gave a 
spectacular increase in accuracy was the elimination 
of rough and irregular welding at the nozzle exit 
circle. 44 

Other factors which were investigated for possible 
effect on accuracy were oscillation of the liquid 
filler in the heads, variation in filler density, and 
launchers with varying lengths of overhead guides. 
On none of these tests were any definite positive 
results obtained. 

18.4.2 Designation and Types 

The older models, called CWB or CWR in CIT 
reports, had no service designations. The CWR-N 
motor is designated 3.25-in. Rocket Motor Mk 5 
Mod 0. Two 7.2-in. heads have been standardized: 
the Mk 7 for chemical fillers and the Mk 9 for TNT, 
the latter being nearly 3 in. shorter to accommodate 
the higher-density filler without increasing overall 
weight. Complete round designations are “Rocket, 
Chemical, 7.2-in., T21” and “Rocket, HE, 7.2-in., 
T24.” 

18,4,8 Launchers and Service Use 

The standard CWR launcher is the CIT 7.2-in. 
Type 2, designated by the Army as “Launcher, 
Rocket, Multiple Artillery, 7.2-in., T32.” It is a 
24-rail launcher 10 ft long, very similar in design to 
the BR crate, which can be mounted on the ground 
or in the bed of a 23^-ton truck as shown in Figure 5. 
Although Army Ordnance produced a considerable 
quantity of the launchers, no service use of the 
CWR is known. 

1844 Reports 

The early development of the CWR is recounted 
in detail in reference 46. Propellant development is 
discussed in reference 47. See also reference 70 for a 
report on high-speed water tunnel tests. 

185 TARGET ROCKETS 

The development of rockets as targets for anti- 
aircraft training antedates the CIT contract. It 
was undertaken jointly by Sections H and E, Divi- 
sion A, NDRC, in August 1941, and three flight 


TARGET ROCKETS 


159 


tests were conducted in the East before OEMsr-418 
was organized. In the summer of 1941, when 
NDRC had circularized the Armed Services as to 
their interest in a variety of proposed rocket pro- 
jectiles and devices, the Coast Artillery, which had 
the responsibility for training antiaircraft gunners, 
answered that they would like to have a target 
rocket developed. Something was needed which 
would give gunners adequate practice in firing at 
targets which approximated the speed and courses 
of aircraft. Small radio-controlled airplanes or 
drones would have served the purpose, but they 
were not in quantity production and were neither 
cheap enough nor fast enough. The conventional 
towed sleeve target was too slow and moved on too 
steady a course to give the necessary training in 
“leading’ ’ a high-speed, maneuvering target. Tar- 
get rockets could be fired toward, away from, or 
across the line of fire of the guns either in low 
straight paths or in a high looping trajectory, and 
should be able to simulate most of the situations met 
in battle. 

The newly formed CIT group took over the de- 
velopment begun in the East. There were several 
advantages in beginning with the target rocket. It 
was a less complicated problem than most others, 
since neither head nor fuze were required. The 
main requisites were velocity and visibility — a motor 
with sufficient thrust and fins of sufficient size . The 
experience in designing the motor and firing the 
target rockets would give useful data which could be 
applied to the more difficult problems which were 
being undertaken while the target rocket develop- 
ment was proceeding. 

The work was handicapped by troubles with pro- 
pellant — both quantity and quality, and the early 
history of the target rocket is the history of the 
development of satisfactory propellant. 48 Rockets 
were made, however, even though for a few months 
propellant failures were rather frequent, and on 
November 29, 1941, the 78th Coast Artillery at the 
Mojave Antiaircraft Artillery Range got a chance 
to shoot at three target rockets. This was a small 
beginning, but “rocket shoots” rapidly became 
bigger and more frequent. The verdict of officers 
and men was uniformly favorable. By the summer 
of 1943, for example, the target rocket range at 
Camp Pendleton, on the southern California coast, 
was scheduled for four weeks in advance. By the 
following December, when the Bureau of Ordnance 
standardized and undertook the production of two 


types of rockets and several launchers, the CIT 
group, developing and producing its own target 
rockets, had participated in training some 21,000 
men. Improvements in the rocket design continued 
to be made up until that time . 

18 5 1 Design Features 

Motor. The first CIT target rocket motors were 
very similar to those which had been tested in the 
East, 3.25 in. in diameter and approximately 6 ft 
long. The only propellant available was 1.0x0.25- 
or 0.87 x 0.25-in. tubes 5 in. long, and they were 
strung on a steel “cage” attached at the front end 
of the motor. In one design, the whole rocket was 
motor, and the gas from the propellant charge at 
the front had to traverse a long empty space to 
reach the nozzle. In another design, the rocket 
was jointed in the middle, the rear part being 
motor and the front part empty tube. The latter 
design was tried out because it was thought that 
smaller heat losses might give better efficiency. 
Tests showed that the long motor with the dead 
space gave slightly smaller gas velocities, longer 
burning times, lower average pressures, and con- 
siderably smaller differentials between peak and 
average pressure. It was chosen as standard pri- 
marily because it was cheaper and easier to produce. 
Fourteen-gauge tubing (0.083-in. wall) was used for 
the motor in order to save weight, and the pressure 
was kept low by using nozzle K ’ s around 180 or 
lower (see Section 22.4). 

Earliest models had machined screw-in nozzles, 
but the spun integral nozzle (see Figure 4 A of 
Chapter 23) soon became standard. A front closure 
involving an obturator cup and a piston ring was 
worked out, so that no threading was required on 
the motor tube. For attaching the fins, L-shaped or 
T-shaped lugs were welded to the tube. 

When single large grains became available, the 
propellant was moved back to the rear of the motor, 
and the grid was seated on a grid ring which was 
held in place by welding through four holes drilled 
through the tubing. 

After set-back fins were developed, the motor was 
simply shortened to 32 in. without other modifica- 
tion. (See Figure 6.) 

Tests of multinozzle motors with the set-back 
fins were made, and three-nozzle motors were found 
to be satisfactory provided that the nozzles were 


160 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


aligned with the fin longerons. Random orientation 
of the nozzles with respect to the longerons appar- 
ently increased the dispersion. 

For its production, the Bureau of Ordnance pre- 
ferred a motor design similar to that of other 
motors in production. Hence 11 -gauge tubing was 


Propellant problems for the target rocket were 
straightforward once the technique of making good 
grains had been learned, and no special comment on 
them is required. 

Fins. The fin problem for the target rocket is 
obviously entirely different from that for any other 



Figure 6. CIT target rockets. Left: eary 4-fin variety. Right: set-back fin variety with straight fins (final 
design has fins canted slightly to impart slow rotation). 


specified and a nozzle design like that of the Mk 7 
AR motor (see Figure 4C of Chapter 23). Stud- 
welded bolts with washers and nuts were also speci- 
fied in place of the L-shaped lugs, and a threaded 
front closure was substituted for the piston ring. 
These changes increased the weight and decreased 
the velocity somewhat, since the propellant grains 
were not changed. 


rocket. In addition to their being as large as possible, 
the chief requirements are that they be relatively 
cheap to manufacture and able to withstand the 
weather. As with any rocket part, it is desirable 
also that they be light. A large variety of fin con- 
structions was tried, and it would serve no purpose 
to detail them here. In all cases they were made 
relatively thick, with an internal structure to with- 


LITTLE ROCKETS 


161 


stand the acceleration and wind forces, and were 
covered with some thin material; aluminum, bur- 
lap, cloth of various types, plastic, and paper were 
among the things tried. 

The 6-ft motors had 4 fins approximately 1 ft 
wide and 3 ft long. After static firing had showed 
that the hot part of the jet does not spread very 
rapidly on leaving the nozzle, tests began on target 
rockets having the fins extending back beyond the 
nozzle exit plane. It was found possible by attach- 
ing them to tapered longerons to set the entire fin 
back of the nozzle and thus do away with the neces- 
sity for a long motor which had been required to 
move the center of mass forward. Improved fin 
construction together with the much smaller accel- 
erations given by the thick-web single-grain charges 
made possible an increase in fin width to 18 in. 
With the wider fin, the number could be reduced 
from 4 to 3 and still increase the visible area over 
that of the long 4-fin target. 

The fin construction which was finally worked out 
as the most satisfactory was to cover the framework 
with bfg-in. fiberboard attached with staples and 
glue, and then to spray the whole fin with a coating 
of hot paraffin to make it watertight. The frame- 
work is made of high-quality white pine with 
tongue-and-groove joints nailed and glued. Such 
fins could be manufactured for less than $7 each 
in small quantities. 

It was found that canting the fins slightly gave a 
considerable decrease in dispersion, and with the 
set-back design it is easy to do. The longerons are 
fastened parallel to the motor axis and the fins are 
attached to them with the rear end displaced % in. 
from the front end. The resulting rotational velocity 
is between 1 and 2 rps. 

Contacts. Electrical contacts on a molded bake- 
lite cap which sealed the nozzle were used for a time. 
Since rapid loading is less important than certain 
contact, they were later replaced by ordinary house- 
hold-type electric plugs which fit into receptacles on 
the launchers. 

1852 Launchers 

The older type target rockets were fired from a 
simple launcher consisting of two pieces of lj^-in. 
tubing with one fin extending down between them. 
With the standard design, this system was not 
practicable, and a so-called “M-rail” was designed 


which contacts the motor tube and two adjacent 
longerons. It can be mounted on the standard Ml 
trailer launcher for Army target rockets. Separate 
trailer mounts and tripod mounts have also been 
used. 49 

18.5.3 Designations and Types 

Target rockets of various velocities from 450 mph 
down to 200 mph were used at various times. They 
were usually differentiated in CIT reports by draw- 
ing number, 3T4, 3T7, etc. Two velocities were 
finally chosen, 290 mph and 415 mph, to be used 
respectively for beginning and advanced training. 
In the final designation, all fast rockets are called 
3.25-in. Rocket Target Mk 1 Mod 0, and all slow 
ones are 3.25-in. Rocket Target Mk 2 Mod 0, except 
that in either case the addition of a flare for night 
firing changes the Mod number to 1. The earlier 
designation distinguished between CIT (Mk 1 fast, 
Mk 2 slow) and Bureau of Ordnance (Mk 3 fast, 
Mk 4 slow) production. 


1854 Reports 

Summary reports which discuss the development 
of the rockets themselves as well as the training 
tactics are references 50, 51, 52, and 53. Reports 
on methods of training, scoring equipment, etc., 
include references 54, 55, and 56. On manufactur- 
ing problems, see reference 57. 


i8 6 LITTLE ROCKETS 

The various rocket projectiles with 1 .25-in. motors 
developed by CIT are so closely related that they 
will be discussed together. They will be treated 
somewhat more briefly than the larger rockets be- 
cause they are less important from the standpoint 
of service use and also because, once the funda- 
mental principles of rocket design had become un- 
derstood, their development was relatively straight- 
forward and simple. They did have an important 
place, however, both in providing information nec- 
essary for the design of larger rockets and in training 
Service personnel in their use. 


162 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 


18 6 1 Chemical Warfare 

Grenade [CWG] 

The chemical warfare grenade [CWG] was one of 
the earliest CIT projects, the first model having 
been fired in the desert test of December 23, 1941, 
which saw also the introduction of its big brother 
the CWR. The specifications laid down by the 
Chemical Warfare Service were payload, 1 lb of 
liquid in a frangible container; range, 600 yd; 
accuracy, 5 mils; trajectory, not to exceed 30-ft 
height in 200-yd range; projector, to be carried and 
operated by one man. Since it was thought of as 
primarily for use against tanks, the rocket was 
originally called the antitank grenade [ATG], but 
this name became obsolete within a month. 

The service history of the CWG was disappoint- 
ing. When it was demonstrated to the Chemical 
Warfare Service in April, the reception was unen- 
thusiastic because the bakelite head could not be 
used for FS, although the observers were pleased 
with its accuracy. Later tests showed that 1 lb of 
liquid was not sufficient to cause the damage re- 
quired. Research on the CWG was therefore 
stopped, but it had served an important function 
in making possible a large number of experimental 
tests with little expenditure of propellant, which 
was extremely scarce, and had yielded much in- 
formation. 

Motor Design Features 

The first motors were necessarily designed to use 
the only solventless-extruded ballistite tube then 
available (}%> in. OD by 34 in • ID) since tests had 
already demonstrated the marked superiority of 
this material to solvent-extruded tube. They were 
made of 1.25-in. outside diameter, 16-gauge steel 
tubing, threaded on the outside to take the front 
closure and on the inside to take the machined 
nozzle. Several motors burst at the undercut 
of the nozzle thread when the motor became hot, 
and the screw-in nozzle design was therefore 
abandoned in favor of the spun integral design. 
Several hundred rounds of CWG motors were 
fired on the yaw machine, and the following facts 
were learned: 

1. Longer nozzles tended to give smaller side 
forces (range firings appeared to corroborate this 
with smaller dispersion) . 


2. Integral nozzles were frequently distorted 
and cocked out of line by the heat and pressure of 
firing. 

3. If the burning time were short enough, the 
distortion did not occur. 

4. Side forces could be reduced by careful align- 
ment of the nozzle exit cone. 

Also shortly after the first CWG firing, the first 
calculations of malalignment effect were made, 
which showed that accuracy could be considerably 
improved by decreasing the burning time. 

On the basis of all this information, a double-web 
charge was designed to reduce the burning time as 
much as possible. It consisted of a 34x34“i n * 
tube inside a 1.0xJ4-in. lube and gave a burning 
time of only 0.12 second at 70 F. Rounds were 
checked on a malalignment balance and carefully 
straightened to keep their malalignments small. 
The result was that the dispersion dropped from its 
original very large value to approximately the 5 mils 
desired. Experiments continued to attempt to 
reduce the dispersion of the single-web charge, but 
they were doomed to failure by the gas malalign- 
ment, which at that time, of course, was not under- 
stood. A large number of field tests of the CWG 
with various launcher lengths, various burning 
times, and various fin sizes during 1942 established 
the correctness of the theory of dispersion which was 
published in reference 58 and formed the basis for 
all subsequent finner development. 

To avoid the complicated assembly operations 
involved with the tubular double-web charge with 
its numerous celluloid strips, a 1-in. 4-spoke or 
“okra” grain was extruded. Since its burning time 
was short, it gave good dispersion. Its gas velocity 
was somewhat smaller because of the slivers left at 
the end of burning, and considerable difficulty was 
found in trying to make the cylindrical portion and 
the spokes burn at the same rate. Interest in this 
grain shape declined with the abandonment of the 
CWG. 

The grid originally used for the tubular double- 
web charge was a complicated structure of four legs 
and a ring. An adaptation of the box grid having 
six pieces instead of four was found to work better 
for both double-web and 4-spoke and was adopted 
as the final standard. 

The double-web CWG is shown in Figure 7. 
Features other than the motor require little com- 
ment. The head was formed of an 18-in. bakelite 


LITTLE ROCKETS 


163 


tube, 1.62 in. in outside diameter, closed at the 
front end by a hemispherical steel cap and extend- 
ing back several inches over the motor to a support- 
ing shoulder of sufficient width to maintain accurate 
alignment. Several types of stabilizing fins were 
tested, including a folding design with four 4x 
1.12-in. blades which opened by means of springs 
from a rectangle 1.62 in. square to a radial position. 
They worked satisfactorily, but were abandoned as 
too complicated, and four fixed radial fins were 
adopted. 

Reports on the CWG 
See references 58, 59, 60, 61, and 62. 


Figure 7. Assembled CWG and cutaway showing 
charge. 

18 2 Subcaliber Rockets 

The acute shortage of ballistite which plagued us 
constantly during the first two years of World 
War II made the development of subcaliber prac- 
tice rounds imperative, although they would have 
been useful in any case. Development of the sub- 
caliber ASR (175 fps) began almost simultaneously 
with that of the drift signal rockets (see Section 
19.1.6), and by adjusting the head weight it was 
possible to use the fast drift signal motor with the 
addition of a ring tail for the practice ASR. 

The increase in velocity to match the 200-fps 
VAR was accomplished without changing the motor 
simply by shortening and lightening the head. It 
was hoped that the 300-fps subcaliber VAR could 


use the same head as the 200-fps, but the very long 
grains gave such low gas velocities that the required 
velocity was not reached. Rather than extrude a 
thicker-web grain which could have given enough 
impulse, it was decided to retain the 1.0 x 0.5-in. 
shape and use another still lighter head. 

The heads are identical except for length and are 
machined from 2.5-in. steel bar stock. Originally 
they contained a shotgun shell which, on contact 
with the target, was set off by a firing pin which 
projected in front of the head. The shells were found 
to be unsafe to handle and unnecessary since the 
thud of the head itself on the submarine hull was 
enough to tell the occupants that a hit had been 
scored. The shell was therefore abandoned. 


nozzle and head construction and double-web tubular 

Launchers for the subcaliber rounds are boxlike, 
formed by bending steel sheet, and adapted to be 
attached to the launchers for the full-scale rounds. 
The one for aircraft firing (2.5-in. Rocket Launcher 
Subcaliber Mk 3 Mod 0) is shown in Figure 8. 

Reports 

The only official CIT report on the subcaliber 
ASR projectile itself is reference 63, which was 
written very early in its development when it still 
used the CWG motor. The subcaliber VAR’s are 
discussed in several of the reports on full-scale 
ammunition (see Section 19.1.4). Reports by the 
Morris Dam group on the underwater performance 
of the various models include references 64, 65, 66, 
67, 68, and 69. 




164 


DESIGNS OF FIN-STABILIZED ROCKETS FOR SURFACE WARFARE 



Figure 8. Subcaliber VAR being loaded into 2.5-in. Launcher Mk 3 mounted on 7.2-in. Launcher Mk 1 Mod 
1 under wing of PBY-5. Compare with Figure 1 of Chapter 17. 


18 ‘ 6 ' 3 Rocket Grenade 

The last rocket having a 1 .25-in. motor which was 
developed was the incendiary rocket grenade [IRG]. 
Its purpose was very similar to that of the CWG, 
which it resembled also in its history, since it was 
tested extensively and brought to completion at just 
about the time that the Services decided that they 
had no use for it. Like the CWG it was to be fired 
at a low angle and required high accuracy, but its 
head was provided with a very simple shear-wire 
impact fuze to disperse the contents rather than re- 
lying simply on the shattering of the head on 
impact. 

The motor was essentially the same as 1.25-in. 


Mks 1 and 2 except that to increase the accuracy a 
nozzle with a long entrance cone and a smaller 
throat was used. The long nozzle gave slightly 
less gas malalignment in addition to lengthening the 
round. The smaller nozzle gave a pressure of about 
2,000 psi at 70 F instead of 1,000 psi for the sub- 
caliber motors, thus reducing the burning time and 
the dispersion. Approximately 5-mil accuracy was 
attained at medium temperatures with a launcher 
only 3.5 ft long. 

Since the rocket was intended to be fired from 
a back-pack launcher which could be worn by an 
infantryman, several rather unorthodox launcher 
designs were tried, but development was not 
completed. 


Chapter 19 


SERVICE DESIGNS OF FIN-STABILIZED ROCKETS 
FOR AIRCRAFT ARMAMENT 

By C. W. Snyder 


191 RETRO ROCKETS [VAR] 

H ardly had the antisubmarine rocket [ASR] 
program (see Section 18.2) gotten under way 
in the summer of 1942 when high-priority experi- 
mental work began on the problem of adapting the 
rocket to aircraft use. The development of the 
magnetic airborne detector [MAD] had made it 
possible to detect a submerged submarine directly 
beneath the airplane, but, by the time the target 
was detected, it was already too late to use the con- 
ventional type of bomb. It was suggested that by 
rocket propulsion a bomb could be given a velocity 
equal and opposite to that of the aircraft so that it 
would fall almost vertically from the point of firing 
and hence could be triggered by the signal from the 
MAD. 

The first question to be settled was which direc- 
tion to point the rocket. During burning while the 
rocket is picking up speed, its velocity relative to 
the ground is less than that of the airplane so that it 
is moving in the same direction as the airplane but 
with decreasing velocity. (See typical trajectory in 
Figure 1.) If the rocket is accelerated by its own 
motor, it will be moving backwards through the 
air during the whole burning time, and in this 
attitude the tail fins will increase the yaw instead of 
decreasing it. It was thought, therefore, that the 
rocket’s flight might be better if it were pointed in 
the direction of the airplane’s motion and pushed 
out of the launcher by an auxiliary rocket, called a 
“mule,” which separated from it after the end of 
burning. 

The first firing of an American rocket from air- 
craft in flight took place on July 3, 1942, when 
several ASR’s were fired backwards from a PBY-5A. 
The “mule” tests were made with a nonstandard 
type of ASR with a round nose and a streamlined 
afterbody, but, for the tests in which the ASR was 
accelerated by its own motor, standard ammunition 
was used. It was quickly found that the latter 
system gave only about one-third the dispersion of 
the former, so the use of the “mule” was abandoned. 


Designation and Types 

The ASR was satisfactory for vertical bombing 
from the PBY, but it was too slow for most other 
airplanes on which installations were contemplated , 
and a new series of motors had to be designed. 
Initial tests were carried out with 2.75-in. motors, 
and a tubular three-ridge grain suitable for this 
caliber was extruded and tested. The development 
was not completed, however, and 3.25-in. motors 
were chosen for the purpose because (1) the grain 
which could be put into a 2.75-in. motor would 
not give the 400-fps velocity required for the B-24 
airplane without unduly increasing the burning 
time, and (2) 3.25-in. motors had already been 
developed for other purposes (it was in fact one 
of the first sizes on which work had been done by the 
project) and it was felt desirable to keep the number 
of different motor calibers to a minimum in the 
interest of standardization and simplicity. This 
decision was made in December 1942, and by May 
1943, three 3.25-in. motors had been developed and 
standardized. 

The terms “retro” and “vertical” have been used 
rather loosely and usually interchangeably to de- 
scribe rockets fired backwards from aircraft , although 
it was originally suggested 1 that “vertical” be used 
to describe bombing in which the rocket’s velocity 
simply canceled that of the aircraft and “retro” 
be reserved to apply to the case where the rocket 
has considerably more velocity than the aircraft so 
that its fall relative to the earth is no longer approx- 
imately vertical. Since the original intention was to 
use truly vertical rather than retro bombing, the 
series of rockets designed for this purpose were 
known as vertical antisubmarine bombs [VAB or 
YASB] and later as vertical antisubmarine rockets 
[VAR] (see Figure 2). In cases where the Mark 
numbers (of which there are two sets as for the 
ASR) are not given, the members of the VAR 
series are usually distinguished by their velocity or 
by their drawing numbers: 7V11, 7V12, and 7V13. 

The three VAR motors differ from each other 


165 


166 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


only in the following respects: (1) motor length, 
(2) grain (originally it was intended to have the 
grains differ only in length, but it turned out to be 
preferable to use a slightly thinner web for the 
shortest one), (3) nozzle diameter and contour, and 
(4) length of igniter leads . The motors are designated: 

3.25-in. Rocket Motor Mk 1 Mod 0 
(7V11, 210 fps); 


3.25-in. Rocket Motor Mk 2 Mod 0 
(7V12, 310 fps); 

3.25-in. Rocket Motor Mk 3 Mod 0 
(7V13, 400 fps). 

All use the Torpex-filled Mk 6 head. Complete 
round designations, old and new, are given in Bal- 
listic Data, 2 which also lists several other com- 
binations. 



Figure 1 . Typical vertical-bombing trajectory. 


RETRO ROCKETS [VAR] 


167 


191 2 Design Features 

Heads. The Mk 6 head is patterned after that 
of the ASR but is slightly shorter and has a thinner 
wall. 

Fuzes. For submerged submarines, the ASR 
fuzes — either HIR or underwater-vane-arming — 
worked satisfactorily in vertical bombing, but it was 
felt desirable to have a fuze which would function 
also against surfaced subs, and development of such 
a fuze was undertaken. Considerable work was done 
with AIR-type fuzes, but the velocity of the rockets 
relative to the air was so low during most of their 


Because of the swaged tube, it is necessary to insert 
the nozzle from the front end, and to obviate having 
to press the nozzle in the whole distance, thus gall- 
ing the inside of the tube, the tube’s inside diameter 
is reduced from 3.01 to 2.9 in. in the region where 
the front end of the nozzle and the grid are located. 
This reduction in port area increases the internal K 
of the motor, but with low-performance motors it is 
not serious. Box grids are used. 

Tails. The tail design is identical with that of 
the final ASR, but the shroud rings are 7.2 in. in 
diameter instead of 7.0, so that the same lug band 
will fit both tail and head. 



Figure 2. Section drawing of one of the VAR series. Others differ only in length of motor tube and nozzle 
diameter. 


flight time and their yaws were so large that it was 
difficult to make a propeller work reliably. The solu- 
tion was the SIR (see Chapter 16) which was armed 
after a specified number of rotations of a flywheel 
by a clockspring. It was designated the Mk 139 
Mod 0 and was used on all vertical bombing rockets 
except the ASR. 

Motors. The relatively low stability of the bar- 
rage rocket [BR] with its 2.25-in. motor and 4.5-in. 
tail had indicated that a considerable decrease in 
tail efficiency could be expected with the 7.0-in. tail 
if the motor tube diameter was made too large. 
Firings of the 2.75-in. motors gave good results, but 
those with 3.25-in. motors indicated decreased sta- 
bility. Hence it was decided to reduce the tube 
diameter to 2.75 in. at the rear, as shown in Figure 
4D of Chapter 23, in order to get more air through 
the tail. Formed insert nozzles were chosen for 
cheapness since accuracy was no problem, the dis- 
persion of the rockets in vertical bombing being less 
than the uncertainty in the position of the aircraft. 


Grains. To minimize the forward travel of the 
rockets, a short burning time was desired, and after 
static tests a web thickness approximately the same 
as that of the ASR grain was settled upon. 

191-3 Launchers and Service Use 

The first and largest launcher installation de- 
veloped for VAR’s was that on the PBY-5A or 
Catalina flying boat. It consisted of 24 channels, 
12 under each wing, formed from Y%-m. Dural sheet, 
the individual rails being fanned outward by vary- 
ing amounts to spread the impact pattern. The 
rockets were hung under the channels on lug bands 
which rode on the turned-in edges of the channels. 
An intervalometer was inserted into the firing cir- 
cuit so that the rockets were fired in three sym- 
metrical groups of eight to give an impact pattern 
approximately 140 ft wide and 40 ft along the 
direction of flight. 

The next plane equipped was the B-18, which 


168 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


carried 16 steel launchers similar to those on the 
PBY. The original launcher design for this airplane 
was a Dural tube in which the rockets fitted fairly 
snugly without requiring lug bands. These were 
objectionable because they caused much more drag 
on the airplane than the flatter channels but prin- 
cipally because the aspirator effect of the nozzle 
reduced the air pressure behind the rocket so much 
that the burnt velocity was reduced as much as 20 
per cent in some cases. The B-24 was also equipped 
with launchers on the fuselage. Neither of the 
Army planes took the VAR into combat, however, 
because sole responsibility for aerial antisubmarine 
warfare was soon assigned to the Navy. 

The PBY and the TBF (carrying 4 launchers on 
the fuselage) actually used the rockets against 
enemy submarines with good effect. Vertical bomb- 
ing proved to be much less useful than had been 
expected, however, because of a change in German 
submarine tactics. Vertical bombing theory assumed 
that the submarine would be submerged or getting 
there as rapidly as possible, but, during the latter 
part of 1943, German subs began staying on the 
surface and fighting it out with their deck guns. In 
this situation it was too dangerous for the attacking 
plane to make a straight run at low altitude as was 
required for vertical bombing. Only in special areas, 
such as the Straits of Gibraltar, where submerged 
submarines attempted to slip into or out of the 
Mediterranean, were the potentialities of the VAlt- 
MAD combination fully realized. 

As development of various installations pro- 
ceeded, the emphasis shifted gradually from vertical 
to retro bombing. Firing backwards with a velocity 
considerably exceeding that of the plane had the 
advantages that (1) the rocket was more stable so 
that yaws on striking the water were smaller, (2) the 
launchers could be pointed slightly downward or 
even at a considerable angle so that flight time and 
hence dispersion could be reduced, and (3) firing 
could be delayed until the plane was somewhat past 
the target, thus simplifying the sighting problem. 
Considerable experimental work was done with the 
BR from the A-20 and the B— 18, but no service 
requirements for the installations materialized. 
Photographs and brief discussions of the installa- 
tions are contained in the summary reports on 
retro bombing, 1,3 but further details can be found 
only in the weekly progress reports of the period. 

Tests of retro firing of 100-lb bombs propelled by 
six ASR motors were also carried out. 4 


1914 Reports 

References 3, 5, 6, and 7 give a complete account 
of the progress of the vertical-bombing program 
from beginning to end, discussing both the ammu- 
nition and the installations. A bibliography of 
various other reports pertaining to particular in- 
stallations is given in reference 3 . 

1915 Related Rockets 

Although the VAR motors did not find extensive 
use in the application for which they were originally 
designed, they were adapted by Army and Navy 
Ordnance for various other purposes. An example of 
this is the “Cutteroo Grapnel,” to propel several 
multipronged hooks and a steel cable. When the 
hooks were pulled back, they could clear out barbed 
wire, detonate land mines, or do other jobs. A 
similar use of the motors, in which CIT was directly 
involved, was in obtaining samples of the earth 
from the center of the crater at the first atomic 
bomb test in New Mexico, July 16, 1945. The 
existence of three motors differing in thrust but 
being otherwise interchangeable made the develop- 
ment of such uses relatively simple . 

The first model of the 3.25-in. Aircraft Rocket 
Motor, the Mk 6, utilized the same nozzle design as 
the VAR’s but with a different tail. It is discussed 
more fully in Section 19.2.2. 

The first “window” rocket (3.25-in. Rocket Mk 4 
Mod 1) used the Mk 2 VAR motor intact except 
for the tail, which was cut down from the tail of the 
Mk 6 motor to enable firing from T-slot launchers. 
The only serious design problems in connection with 
the window rocket had to do with the ejection 
charge in the head, which has been discussed in 
Chapter 16. The purpose of the rocket was to eject 
at the peak of its trajectory its load of metalized 
paper strips to confuse enemy radar. Although the 
range of the rocket was rather short and the relia- 
bility of the ejection charge not all that could be 
desired, it was effectively used at the time of the 
Normandy landings and later. 

As discussed in Section 18.4, the final model of 
the chemical warfare rocket [CWR] used the Mk 3 
VAR motor with a thicker- web charge. 

The Mk 3 motor with a special tail was used also 
for the smoke float rocket, which was developed in 
1943. 8,9 


RETRO ROCKETS [YAR] 


169 


19.1.6 Drift Signal Rockets 

As soon as it had been shown that vertical bomb- 
ing of submarines was feasible and the tactics began 
to be worked out, it was apparent that a means of 
marking the submarine's position was required. The 
marker should be rocket-propelled so as to duplicate 
the trajectory of the vertical bombs themselves. 
The patrol plane could then cruise at low altitude 
over the ocean, and, if a magnetic signal like that of 
a submarine was received by the MAD , it would fire 
a marker. By making several passes over the sub 
and releasing a marker at each contact, its course 
and speed could be estimated so that the actual 
bombing attack could be made with the best chance 
of success. 

Standard Navy drift signals were to be used, the 
weight of which was only about 4 lb (the one finally 
adopted weighed only 3 lb), so that sufficient 
velocity could easily be attained with 1.25-in. 
motors. The problem was to get the right velocity 
to match the trajectory of the VAR and to ignite 
the flare heads. The first models utilized the chemical 
warfare grenade [CWG] motor with the 4-spoke 
charge (see Section 18.6.1), and rather extensive 
static-firing tests of various lengths of grain were 
made, since velocities as low as 35 knots were under 
discussion. It was found that the shorter grains 
required higher nozzle K’s for satisfactory burning, 
even as high as 250, and that radial holes greatly 
improved the burning characteristics of the multi- 
web grains just as it had for the tubular. Percussion 
ignition was used on the earliest models, a spring- 
loaded firing pin firing a .32-caliber blank cartridge 
into 4 g of Quickmatch in the front end of the 
motor. This was soon abandoned for electrical 
ignition, however. 

The short burning time given by the 4-spoke 
grain was no advantage for this motor, and it was 
more complicated to extrude. A 1 .0 x 0.5-in. three- 
ridge tubular grain was therefore specified, and it 
became standard for all 1.25-in. motors. Trial of 
several nozzle designs resulted in the adoption of 
the swaged-in machined nozzle shown in Figure 3B 
of Chapter 23, and it, too, became standard. Ig- 
nition of the flare head was easily accomplished by 
letting the thrust of the motor shear a wire, allow- 
ing the motor to move forward in a sleeve and strike 
a blank cartridge in the rear of the head . After the 
thrust ceased, motor and head were separated by a 
spring, since the weight of the motor would other- 


wise sink the head. The burning time of the motor 
was so short and the ammunition dispersion so 
small a fraction of the overall dispersion that no tail 
was found to be required on the motor. 

During the course of bombing experiments at the 
Goldstone Range in the Mojave Desert, one im- 
portant lesson was learned the hard way when a 
motor ejected its nozzle, which pierced the closed 
breech of the launcher, narrowly missed the man 
who was loading it, went out through the wall of the 
cabin, and entered the gasoline tank in the wing of 
the PBY, where it was brought to rest by the high 
drag of the gasoline. After the several hundred 
gallons of precious fuel had drained out and evapo- 
rated, the nozzle was recovered and found to be 
neatly plugged by a little tube of cellulose acetate, 



Figure 3. Drift signal rocket in Mk 2 launcher. 


the squib compartment of the molded plastic igniter 
which had just become standard a few days before. 
Thereafter igniters for small-nozzle rockets were 
carefully designed to avoid any possibility of large 
fragments. 

In CIT reports, these rockets were first called 
vertical flare bombs [VFB] or vertical flare rockets 
[VFR] and later vertical float lights. Earlier Navy 
designation was Drift Signal Rockets Mk 15 and 
Mk 16, but the latest nomenclature drops the Mark 
numbers and specifies them by velocity — 200 fps 
and 300 fps. Other slower models were worked on 
but not standardized. 

Launchers 

The tactics of submarine hunting with aircraft 
required that it be possible to fire a considerable 


170 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


number of drift signals, and a reloadable launcher 
was therefore indicated. A closed-breech launcher 
was designed with a loading door on the side at the 
rear. It projected backward and about 15 degrees 
downward through a hatch in the under side of the 
plane near the tail. The flare head ran on guide rails 
inside the main tube with about 1 in. clearance on 
all sides so that no pressure built up inside the 
launcher. After the accident at Goldstone, the 
breech of the launcher was reinforced with a steel 
plate. A twin launcher of this type (see Figure 3) 
was adopted as standard and designated the Air- 
craft Rocket Launcher Mk 2 Mod 0. 

The drift signal rockets were made exclusively for 
vertical bombing, and their service record is there- 
fore identical with that of the VAR’s. 

Reports 

The drift signal rocket is discussed in most of the 
reports on 7.2-in. retro rockets (see Section 19.1.4). 
See also reference 10. 


192 3.5-IN. AND 5.0-IN. AIRCRAFT 

ROCKETS [AR] 

Development History 

When the CIT group began, the development of 
medium- and high-altitude antiaircraft rockets was 
one of the principal projects assigned to it, because 
the development of such a high-performance rocket 
would necessarily depend on the development of 
techniques for dry-extruding very large propellant 
grains. A few field tests of such rockets were made 
in the early days of the project, but little progress 
could be made until the 8-in. extrusion press was 
put into operation in April 1942, and by this time 
the ASR, BR, and other rockets had taken a higher 
priority. Some work on high-performance motors 
went on during 1942, but it was mainly with 2.25-in. 
motors because grains and metal parts could be 
produced cheaply in this size. 

In the early spring of 1943, with the virtual com- 
pletion of experimental work on the BR, the prob- 
lem of designing a 3.25-in. motor with as large as 
possible a propellant grain was attacked with vigor. 
This caliber was chosen because dies for extruding 
tubular grains were available and because it was 


desired to duplicate with ballistite the performance 
of the British cordite-propelled UP3, which had 
been put into service in 1941. 

During March and April, static and field tests 
were made with CIT-extruded tubular grains of 
cordite and ballistite in 14- and 11 -gauge motors 
and the following results were established: 

1. With either propellant, a refractory coating is 
necessary on the interior of the 14-gauge tubes, 
since otherwise heat failures are experienced at high 
temperatures. 

2. The thickness of the 11 -gauge tubes is suffi- 
cient to make refractory unnecessary with up to 6 .8 
lb of ballistite, but a fairly small increase in burning 
time might make the motor unsafe because of 
heating. 

3. In static firing, grains of 2.5 x 0.4-in. three- 
ridge ballistite weighing 6.2 lb were satisfactory up 
to 130 F with either rod stabilization or radial holes, 
and 6.8 lb was satisfactory with rod stabilization. 
In the field with the addition of the setback force, 
the 6.2-lb rod-stabilized grain was satisfactory at 
high temperatures, but the other two were not. 

In May the Bureau of Ordnance requested devel- 
opment of a ship-to-shore rocket to have a range 
not less than 10,000 yd with any of three inter- 
changeable heads: (1) a light-case head for chemical, 
smoke, or high explosive for blast effect, (2) a high- 
explosive fragmentation head, and (3) an incendiary 
head. The motor was specified as 3.25-in. diameter, 
and, although the head weight was not specified in 
the directive , initial experimentation was conducted 
with a head having the weight of a 75-mm shell, 
approximately 13 lb. It was found that the rod- 
stabilized 6.8-lb grain would achieve the required 
range but the 6.2-lb grain would not. Neither was 
satisfactory, however, because the stabilizing rod 
was eroded through and ejected white-hot near the 
end of burning. 

Meanwhile, when the difficulties involved in 
increasing the weight of a tubular charge had begun 
to become apparent, the propellants section had 
commenced work on extrusion dies for a cruciform 
charge, 11,12 the British having had good success with 
a grain of similar shape. After the technique of 
inhibiting these grains and the proper arrangement 
of inhibiting strips had been worked out, a 9-lb 
cruciform grain gave excellent performance stat- 
ically even at 140 F . This weight was sufficient to 
give 10,000-yd range to a 20-lb head provided that 
the fuze had a small enough drag. 


3.5-IN. AND 5.0-IN. AIRCRAFT ROCKETS [AR] 


171 


Early field tests of the upper temperature limit 
of this grain were complicated by the fact that the 
grid seating surface at the front end of the nozzle 
was too narrow, so that, when motors burst and the 
recovered grids and nozzles gave evidence that the 
grid had slipped, it was impossible to determine 
whether the grid slippage had been the cause or the 
result of the high pressure which burst the tube. 
Numerous failures at both 120 F and 130 F were 
experienced, and even yet the full explanation for 
them is not known. They nearly always happened 
after the motor had left the launcher, but the time 
of burst varied all the way from less than one-third 
burnt up to nearly seven-eighths burnt. Failures 
occurred at the nozzle end, in contrast to the be- 
havior of all other motors, but was not the result of 
heating because the camera records showed that 
they were preceded by a definite increase in acceler- 
ation and in the luminosity of the jet, presumably 
because broken pieces of powder began to be ejected 
and to burn outside the nozzle. One piece of grain 
was recovered with a piece of inhibitor strip attached 
which showed that the front ends of the inhibitor 
strips were completely eroded away before the 
middle of the burning time, and this was imme- 
diately confirmed by partial burnings with 11 -gauge 
motor tubes, previous firings having failed to dis- 
close it because they were made with thicker tubing 
which did not heat up so much. This experience 
taught us another important lesson, that static- 
firing tests should be done with completely standard 
motors. 

Even after the inhibitor strips had been increased 
from 0.05 to 0.10 in. in thickness to prevent their 
eroding away and decreased from 8.5 to 7.5 in. in 
length so as to make the burning more regressive 
and reduce the end pressure peak, and the grid 
seating surfaces had been made adequate, occa- 
sional bursts at 120 F, frequent bursts at 125 F, 
and about 80 per cent bursts at 130 F occurred. An 
increase in nozzle diameter from 1.44 to 1.50 
improved the performance greatly, giving only one 
burst out of 33 rounds at 130 F and none out of 100 
at 120 F. Nevertheless, in proof firing a few weeks 
later, one motor burst at 120 F. This burst impelled 
the decision, which had previously been contem- 
plated, to reduce the grain weight from 9 to 8.5 lb in 
order to raise the upper temperature limit to the 
point where the rockets could be proof-fired at 130 F 
and approved for service use up to 120 F. A some- 
what later change in motor design, which reduced 


the internal K slightly, gave still better high -tem- 
perature performance, as mentioned later. 

The adoption in September 1943, of the 8.5-lb 
grain, later designated the Mk 13 grain, solved the 
most difficult problem of the 3.25-in. AR motor. 
While the propellants and motor design groups had 
been preoccupied with this problem, other develop- 
ments had been taking place which had changed 
the nature of the rocket drastically. The British 
success in adapting the UP3 to antisubmarine use 
from aircraft strongly indicated the desirability of a 
parallel development in this country. Thus in early 
June, a 20-lb solid steel head had been designed, 
and aircraft forward-firing launcher development 
had begun. Forward-firing tests from airplanes in 
flight, first with British rockets and soon with CIT 
rockets, became more and more frequent. By the 
middle of August, the AR had been assigned the 
highest priority among all the antisubmarine weap- 
ons, and CIT had been requested to manufacture 
10,000 rounds per month until Navy contractors 
could get tooled up to begin. One month later, the 
request had been increased to 100,000 rockets in six 
months, and this number was actually delivered by 
the end of the following March. 

Although the original purpose of the 3.5-in. AR, 
that of puncturing holes in submerged submarines, 
required only a solid steel head, other uses of the 
rocket developed much more rapidly than the rocket 
supply, and other heads were designed. In par- 
ticular, the base of the 5.0-in. A A common shell 
was boat-tailed and bored out to take a motor 
adapter and became the 5.0-in. Rocket Head Mk 1. 
The combination of this head with the 3.25-in. 
motor became known as the 5.0-in. AR and ulti- 
mately overshadowed the 3.5-in. AR in importance 
as the submarine menace declined. 


19.2.2 Motor Design Features 

The first motor which was extensively tested was 
the 3 A9 (designated by its drawing number series) . 
Since it was designed for long-range firing, it was 
made as smooth as possible on the exterior. At the 
rear, the motor tube was swaged to a smaller diam- 
eter for a distance of 6 in. to allow the tail, consist- 
ing of four radial fins attached to a cylinder (similar 
to the CWG; see Figure 7 of Chapter 18), to slip 
over it without increasing the outside diameter. 
The head was the same diameter as the motor tube 


172 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


and was attached by screwing into an internally 
threaded ring held in the motor by a piston ring and 
sealed with an obturator cup. The only protuber- 
ances beyond the 3.25-in. diameter were the four 
fins and four little buttons, two at the front and 
two at the rear, which, in addition to supporting 
the round in a T-slot launcher, held the threaded 
ring and the fins in position. The primary diffi- 
culties with this design were that the rather com- 
plicated front closure increased the motor loading 
time and the lug buttons, which were simply 
threaded into place, were not thought to be safe 
enough for aircraft use where constant vibration for 
long periods might loosen one of them. 

When it became apparent that the principal use 
of the rocket would be from aircraft at relatively 
short ranges, where drag was no longer important 
but greater stability was desirable, it was decided 
to make the head 3.5 in. in diameter so that it could 
be threaded onto the motor and to increase the fin 
size from 3x6 to 5x8 in. The fin change neces- 
sitated a redesign of the rear end of the motor, and 
to use available tooling it was made identical with 
the VAR motors which were in production. It was 
thought that it might be desired to fire this motor 
with a VAR head and tail, but it turned out that 
this combination had insufficient stability and was 
extremely wild. The new motor was the 3A12 and 
it became the first standard service AR motor, 
designated Mk 6. With a solid steel head (3.5-in. 
Mk 1), it formed the 3.5-in. AR Model 1 (see Chap- 
ter 17, Figure 2). 

The 3 A 1 2 was soon abandoned because the manu- 
facturers which the Bureau of Ordnance chose to 
produce the rocket in quantity objected to the 
complicated shape of the nozzle end of the motor, 
and the 3A16 or Mk 7 motor was designed in close 
collaboration with them to make it as adaptable to 
quantity production as possible. The motor design 
involving the bead at the front end of the nozzle is 
discussed in Chapter 23 and is shown in Figure 4C 
of Chapter 23. It had an important advantage from 
the standpoint of ballistics also in that eliminating 
the swaged portion ahead of the grid decreased the 
internal K slightly.® Since the primary difficulty 
with the upper temperature performance of the 
motor was its high internal K, this was expected to 
alleviate the difficulty. An interesting analysis of 

n Internal K is the ratio of the burning area of the grain to 
the port area around the grain through which the gas must 
escape. See Section 22.4.2. 


the actual effect of the change is given in a weekly 
progress report. 13 The data must be qualified with 
the statement that the number of rounds fired at 
the extremely high temperatures was not sufficient 
to give a low probable error, but, taken at their face 



Figure 4. 5.0-in. AR Model 12 as fired from T- 

slot launchers and 3.5-in. AR Model 5 as fired from 
post launchers. 

value, they show that the probability of a motor 
burst with either motor increases very rapidly above 
140 F but is less than half as great for the 3A16 
as for the 3A12. 

The Mk 7 design was found to be quite satis- 
factory, and production of the motor ran into the 


3.5-IN. AND 5.0-IN. AIRCRAFT ROCKETS [AR] 


173 


millions, CIT contributing the first one-tenth mil- 
lion. The first part of CIT’s production necessarily 
used welded tubing, and, despite pressure tests on 
the motor, occasional bursts during firing occurred 
along the weld line. For its own and the latter part 
of CIT’s production, the Navy procured seamless 
tubing of NE 8735 steel, which had more than 
adequate strength and eliminated this difficulty. 
This considerable increase in tubing strength left 
the nozzle the weakest portion of the motor, with 
the result that occasional nozzle failures were expe- 
rienced in Navy proof firing at high temperature. 
There were three possible causes of the failures: 
(1) The nozzle exit cone may have had a very thin 
spot on one side so that the gas pressure in the 
annular space between it and the nozzle bulged it 
inward. (2) The end of the nozzle skirt may not 
have been brazed securely to the tube in one section 
so that again the pressure could bulge it inward and 
tear it away from the tube . (3) The grain may have 
been faulty so that the motor pressure simply rose 
to such a value that the weakest point had to yield 
even though it may not have been faulty. To the 
writer’s knowledge, it was never determined which 
was the cause, although it was the opinion of the 
projectile section at CIT that (3) was the most 
likely. 

A multinozzle design for the AR motor was exten- 
sively tested and laid the groundwork for the design 
of the HYAR motor. Six nozzles in a circle and a 
central blowout nozzle were machined in a steel 
nozzle plate which was threaded into the motor 
tube. Carefully made motors of this type gave a 
dispersion of approximately 15 mils, which is a 
considerable improvement on the standard model. 
In reply to a request from the Bureau of Ordnance 
for a nozzle design that would eliminate the failures 
occurring in proof firing, the multinozzle design was 
recommended by CIT, but it was never put into 
production. 

Grids. The obvious shape for a grid for a cruci- 
form grain was cruciform, and once the thickness of 
the four arms (H in. wide and % in. thick) had been 
determined by a few static tests, no further change 
in grid design was made except to reword the speci- 
fications slightly whenever anyone thought up a new 
and better method for manufacturing them. The 
only difficult grid problem was how to hold the grain 
on it so that it would not rotate. The earliest 
method was to rivet a celluloid end washer to the 
grid and cement the grain to the washer with cellu- 


solve. An immediate improvement on this design 
was to cement a second washer to the grain and then 
cement the two washers together, thus protecting 
the rivet heads from any possibility of erosion. The 
indexing pins (see Figures 15 and 16 of Chapter 22) 
were soon adopted as simpler and perhaps surer, 
although there was no evidence of failure of the 
other system. 

The so-called “button grid,” a design which would 
make it unnecessary to orient the grain in a partic- 
ular way, was extensively tested. The grain rested 
on a steel disk, 1.38 in. in diameter, supported on 
the nozzle by a spider. The legs of the spider were 
far enough removed from the end of the powder 
grain so that adequate clearance for the gas passage 
was provided even though the spider legs might fall 
directly between the arms of the grain. Static and 
field tests showed only negligible differences in per- 
formance between this and the standard grid, but 
partial burnings showed that, because of the smaller 
bearing area of the button than of the standard grid, 
the end of the grain was deformed around the but- 
ton. Since the difficulty at the upper temperature 
limit was almost certainly due to too great forces 
on the grain, it was believed that the decreased 
support would surely reduce the upper temperature 
limit. Not enough rounds were fired to confirm or 
refute this belief, but it did appear that the effect 
was relatively small. For shorter cruciform grains 
where the forces are not so great, button grids 
appeared very promising and were later used in the 
spinners. 

Lug Bands. The 3A9 motor had the threaded 
button lugs as already mentioned, and the original 
lug bands for the 3A12 were of Dural with a riveted 
button of the same shape. These simple lugs were 
satisfactory because the T-slot launcher was shaped 
so as to bear on the cylindrical portion of the band 
or motor to provide the sway bracing. The fabrica- 
tion of the launcher in this shape was difficult, how- 
ever, so it was decided to make the launcher surface 
smooth and put the sway braces on the lug bands. 
Large numbers of the 3A12 lug bands (so-called 
“3.0-in. lug bands”) were on hand, and they were 
adapted simply by riveting little steel “ears” on the 
Dural band. With the appearance of the 5.0-in. 
head, the 5.0-in. lug band was designed and became 
standard for all motors. The strange shape of the 
clamping mechanism on the 5.0-in. band was in- 
tended to balance the large air drag of the opposite 
side of the band. It did not prove to be a very good 


174 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


design, but the one illustrated in Figure 12 of 
Chapter 23 was not thought of until a year later. 
Since the bands were designed to fit the Mk 4 
launcher which became almost immediately obso- 
lete, they were not very well adapted for the zero- 
length launcher and should not be considered a 
model to be copied. To fit the rear post of the zero- 
length launchers, the “tunnel” lug band was de- 
signed. Its large “ears” serve the purpose of holding 
the tail fins in proper orientation, as does the sway- 
bracing structure of the 5.0-in. band (see Figure 5). 



Figure 5. Front and rear lug bands for 5.0-in. 

AR with 3.25-in. motor as fired from post launch- 
ers. 

Tails. The simplest design of a radial-fin tail is 
to form each fin and one quarter of the cylinder in 
one piece and weld the four identical pieces together, 
and this was the method adopted for all 3.25-in. 
AR motors. Since the motors were light enough to 
be packed four in a box with four fins nested be- 
tween them, the bulk of the tail assembly did not 
appreciably increase the shipping volume. The 


3A12 tail had a threaded ring which screwed onto 
the rear of the motor tube and was held with a set 
screw. The 3A16 design was much more satis- 
factory, being simply slipped onto the tube and 
held by the tail ring which screwed on separately, 
and the 5.0-in. lug band fitting between two adjacent 
fins kept the tail from rotating out of position. The 
primary difficulty with the tail was that the bumped- 
in portions (between the slots which can be seen in 
Figure 4) were not always made the proper depth in 
quantity production so that many tails were exces- 
sively tight and difficult to assemble. It was 
aggravated by the fact that the 8735 tubing tended 
to have a larger diameter than the original tail 
dimensions had contemplated. In addition, the 
kind of handling to which rockets are subject in 
service resulted in the fins being rather frequently 
bent, causing wild dispersion. The double-fin de- 
sign of the HVAR (see Section 19.4.1) was much 
preferred in this regard. 

Electrical Contacts. It was originally expected 
that large numbers of British RP-3 would be used 
in antisubmarine warfare, and it was therefore de- 
sirable to have the two rockets as nearly inter- 
changeable as possible. The British were using a 
large and rather complicated electric plug for 
attaching the igniter leads to the launcher, and a 
simpler die-cast version of it was adopted as stand- 
ard for American aircraft rockets. The plug was not 
very satisfactory; it was bulky, not waterproof, and 
easily damaged, and in addition it took too long to 
attach it to the launcher. Near the end of World 
War II, plans were made to replace it with a smaller 
plug which would avoid the difficulties and which 
would become standard for both Army and Navy 
rockets, but the change had not been accomplished 
when CIT ceased production. 

Caps. Because of the weight of the grain, it was 
deemed desirable to provide more positive support 
for the front end of the grain than was given by the 
fiberboard seal. A die-cast cap was designed, which 
threaded onto the front end of the motor, and in the 
space between it and the seal were inserted a length 
of cardboard tubing and enough perforated card- 
board washers so that the cap would absorb the 
impact of the grain if the motor were dropped on its 
nose. The thermal expansion and contraction of the 
grain was provided for by a thick felt washer in- 
serted between the seal and the igniter. As an addi- 
tional safeguard against moisture, it was desirable 
that the front cap be fairly watertight, but yet it 


3.5-IN. AND 5.0-IN. AIRCRAFT ROCKETS [AR] 


175 


should not hold more than about 50 psi internal 
pressure so that the motor would not be propulsive 
in case of accidental ignition. An attempt was 
made to groove the bottom of the cap so that it 
would blow out at low pressure, but this proved to 
be difficult and to require too close tolerances. The 
bottom was punched out of the cap, and it became 
merely a threaded ring which held a flat steel disk 
and a fiber washer against the end of the motor 
tube. This system was satisfactory. 

On the nozzle end of the motor, a drawn steel 
cup, held in place by the tail ring, also acted as a 
secondary moisture seal and held the electric plug. 
Rendering it nonpropulsive was a more difficult 
problem than for the front cap, and it was finally 
solved by the blowout patch, which later became a 
standard seal component (see Figure 13 of Chap- 
ter 23) . 

1923 Heads 

The first head to be used in service was the 3.5-in. 
Mk 1 (Navy production Mk 2) copied from the 
British head for use against submarines. It is shown 
in Figure 4 of this chapter and in Figure 3 A of 
Chapter 15. It was replaced by the double-ogive 
Mk 8 (see Figure 3C of Chapter 15) after the latter 
had been shown to have a much longer lethal range, 
and in fact the British also adopted it. Two other 
3.5-in. heads deserve mention although their service 
use was, to the best of the author’s knowledge, very 
limited. They are the Mk 6 smoke head (BuOrd 
Mk 9) and the Mk 3 high-explosive head (BuOrd 
Mk 5). Both carried too small a payload (9.4 lb 
of FS or 2.3 lb of TNT) to be very useful. Probably 
more AR’s were fired with the 5.0-in. Mk 1 head 
than with all others. Earliest models had nose fuzes 
only, but the later practice was to supply them with 
a PIR base fuze and a SAP steel nose which could 
be replaced by a nose fuze in the field. By means of 
the fuze-arming solenoid, the rocket could be fired 
so that either the nose fuze or the base fuze func- 
tioned, depending on the type of target. 

19 - 24 Fuzes 

The earliest nose fuze, the Mk 148, was adapted 
from the Mk 137 BR fuze by using a smaller pro- 
peller, a protective cap which was removed when 
the rocket was loaded on the plane, and an adapter 
to fit the threads in the fuze liner. As soon as pro- 


duction could get under way, it was replaced by 
the Mk 149 (see Figure 4) which was specifically 
designed for aircraft rockets and has a streamlined 
body and a waterproof cap assembly which covers 
the propeller and protects the working parts of the 
fuze from weather and icing until it is fired. It has 
also an acceleration-actuated shutter-locking pellet 
which delays the completion of arming until the 
end of burning. The first base fuze, the Mk 146 
with no delay, was later replaced by the Mk 157 
with 0.02-second delay. 

192 5 Launchers and Service Use 

The forward-firing launchers have been described 
in Chapter 17 and their use in service was so exten- 
sive and so well publicized that there is no reason for 
saying much about it here. The first submarine kill 
in which the AR was used was in the Atlantic on 
January 11, 1944. In this and in most subsequent 
submarine attacks, however, it was difficult to 
assess accurately the effect of the rockets because 
machine guns and depth charges were also used 
and because, as one Navy report slyly remarked, 
“the survivors never survive so that they can be 
questioned.” The first use of the 5.0-in. AR in the 
Pacific was in a strike against Rabaul by Marine 
Squadron VMTB-134 which, unexpectedly finding 
itself in possession of 20 sets of Mk 4 launchers, had 
equipped its TBF’s without the aid of any instruc- 
tions, located a shipment of rockets, rescued it from 
the ship’s hold by cutting through a bulkhead 
rather than unload the ship, and then trained them- 
selves for three days. Although theoretically 
rendered obsolete by the HVAR, the 5.0-in. AR 
continued to be used in large quantities in the 
Pacific until the end of World War II because the 
HVAR was not available in sufficient quantity until 
the spring of 1945. It was most successful against 
point targets: A A positions, ammunition dumps, oil 
storage, planes on the ground and in revetments, 
small buildings, and shipping. It was particularly 
effective against shipping, including destroyer es- 
corts, and it is on the record that rockets even sank 
one full-size destroyer. In the Iwo Jima and Oki- 
nawa operations, besides the uses just outlined, 
rocket-firing planes were frequently called on for 
ground support, especially against Japanese caves. 

Surface-fired 5.0-in. AR’s were also used for bar- 
rage where longer ranges than that of the BR were 
required. The T-slot Rocket Launcher Mk 30 


176 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


Mod 0 (essentially identical with the CIT Type 31C 
shipboard launcher shown in Figure 6) saw some 
service in this application, notably on the LSM-R. 
PT boats and LCI’s also fired them from Mk 4 
launchers rigged up in a supporting frame. For the 
Okinawa operation, in addition to other craft with 
automatic 5.0-in. spinner launchers, eight LSM's 



Figure 6. Type 31C shipboard launcher loaded 
with 3.5-in. AR’s with Mk 3 heads. 


were equipped with 480 Mk 4 launchers in close 
array loaded with 5.0-in. All's and were used to 
good effect, although one was put out of action by a 
suicide plane. 

Designations and Types 

With two different motors, five or six different 
heads, and five different fuzes, if the SAP nose is 
included, the nomenclature required to keep all the 
possible combinations straight becomes rather in- 
volved and will not be given in detail. The CIT 
designations most often met in the literature are 

3.5-in. All Model 1 — Mk 6 motor with Mk 1 head. 

3.5-in. All Model 5 — Mk 7 motor with Mk 1 head. 

3.5-in. All Model 14 — Mk 7 motor with Mk 8 
underwater head. 

Latest Navy designation is 3.5-in. Rocket Mk 1 


Mod 0 for all rounds with solid heads, Mk 1 Mod 1 
for all with TNT heads, and Mk 3 Mod 0 or Mod 1 
for all with smoke heads. The Mk 7 motor with the 
5.0-in. head is 5.0-in. Rocket Mk 1 Mod 0 with 
either of the nose fuzes but no base fuze, Mk 1 
Mod 1 with nose fuze and Mk 146 base fuze, Mk 1 
Mod 2 with nose fuze and Mk 157 base fuze, and 
Mk 1 Mods 3 or 4 with no nose fuze but Mk 146 or 
Mk 157 base fuze. A more complete list of designa- 
tions is given in Ballistic Data. 2 

192 7 Reports 

The complete development of the 3.5-in. AR is 
sketched in two CIT reports: references 14 and 15. 
The latter contains a complete index to the weekly 
progress reports and bibliography of reports issued 
up to the time of its publication. Reports on 
underwater tests and trajectories include references 
16, 17, 18, and 19. Pilot production methods are 
discussed in references 20 and 21. The more im- 
portant reports on forward-firing ballistics include 
references 22, 23, and 24. The use of the rockets in 
forward firing is discussed in references 25, 26, 27, 
28, and 29. In addition, there are a large number of 
reports entitled Forward Firing of (Blank) Rockets 
from (Blank) Aircraft, issued by the Army, Navy, 
and CIT. 

193 2.25-IN. AIRCRAFT 
ROCKETS [SCAR] 

The development of the 2.25-in. subcaliber aircraft 
rocket [SCAR], usually pronounced “scar," began in 
January 1944. Since the purpose of the rocket was 
training pilots in firing the larger aircraft rockets, 
it would have been desirable to duplicate the 
standard trajectories exactly. This was realized to 
be impossible, however, since a small rocket would 
necessarily have a considerably larger deceleration 
coefficient because of its small weight, a shorter 
burning time, and a different variation of velocity 
with temperature. The specifications therefore 
called for subcaliber rockets to duplicate as nearly 
as possible the trajectories of the 3.5-in. and 5.0-in. 
All's at 70 F, 20-degree dive angle, 230-knot air- 
speed, and 1 ,000-yd range, these being conditions 
which were frequently used in training. To simplify 
manufacturing, it was desired to use the ASR-BR 
nozzle without modification if possible. 


2.25-IN. AIRCRAFT ROCKETS [SCAR] 


177 


To match the 1,120-fps velocity of the 3.5-in. 
AR in the 2.25-in. caliber even with no payload 
required a high-performance motor unless unusual 
measures were taken to lighten the motor tubing. 
Preliminary calculations indicated that 1.85 lb of 
propellant would be needed. Considerable experi- 
ence with grains in this weight range had already 
been acquired. In the summer of 1942, attempts 
had been made to increase the length of the ASR 
grain above the standard 11.6 in. Tests were made 
on 14- , 16-, and 18-in. lengths, and even on the 
shortest it was found impossible to get satisfactory 
performance above 120 F with the 1.7 x 0.6-in. 
powder. Thicker-web grains, 1.7 x 0.25-in., worked 
better, but on a projectile like the ASR or BR their 
longer burning time would greatly decrease the 
accuracy. Throughout 1943, experiments on 2.25- 
in. motors, usually with thinner walls than 11 
gauge, were carried on to learn about the factors 
which determined the amount of powder which 
could safely be used, and a 2.25-in. rocket, some- 
times called an antiaircraft and sometimes an anti- 
tank rocket, was standardized. 30 With no payload, 
velocities as high as 2,600 fps had been achieved 
with it. 

The restriction to 11 -gauge tubing brought the 
attainable velocity of a 2.25-in. rocket down to the 
neighborhood of that actually required, and, when 
the SCAR was first proposed, there was some 
doubt as to whether a tubular grain could be used 
satisfactorily. The propellant problems were solved 
successfully without recourse to special grain shapes, 
however. 

By April 1944, CIT production of metal parts for 
Navy use was in excess of 1,000 per day and of 
complete loaded rockets in excess of 300 per day. 
The rate of metal parts production soon doubled, 
and total production was more than 200,000. The 
Navy’s own production was, of course, many times 
greater. 

Types and Designations 

The Model 1 SCAR, intended to duplicate the 
3.5-in. AR trajectory, has an overall length of 29.2 
in., of which 26 in. is motor, the head being simply 
a hollow streamlined motor closure. Its grain weighs 
1.75 lb. The motor is Mk 10 or Mk 11 according to 
whether it was produced by CIT or BuOrd and Mod 
0 or Mod 1 according to whether it has a screw-in 
or a formed brazed-in nozzle. Heads are Mk 1 or 


Mk 3. All variations are designated 2.25-in. Rocket 
Mk 1 Mod 0. 

For matching the 5.0-in. AR, the simplest pro- 
cedure at first appeared to be to use the same motor 
with a heavier head, and more than 10,000 of these 
rockets, the CIT Model 2, were manufactured. The 
opposite alternative, using the same head but a dif- 
ferent motor, was soon adopted, however, as the 
Model 3 . It differs from the Model 1 only in having a 
nozzle throat small enough to accommodate its 1 .12- 
lb grain. It was made only in the formed-nozzle 
version, motors Mk 12 and Mk 13. All varieties of 
the slow SCAR are designated 2.25-in. Rocket 
Motor Mk 2 Mod 0. 

19 3 2 Design Features 

Grains. The first calculations indicated that 1.85 
lb of 1.70 x 0.26-in. ballistite would be required to 
give the necessary velocity. This grain gave an 
internal K in excess of 150, so that, as was expected, 
static firing indicated that variations in external 
diameter gave large differences in performance. 
Thus at 130 F, a grain with an external diameter of 
1.71 in. gave a maximum pressure drop along the 
grain of 285 psi, whereas a grain only 0.02 in. 
larger gave 450 psi. If the outside diameter were 
carefully controlled, it appeared from static tests 
that the grain would probably be satisfactory up to 
perhaps 100 F. Effective gas velocities of this 
charge in field firing were higher than expected, so 
that a reduction in charge weight was possible. 
After tests with 1.70 lb, which had too low a nozzle 
K (the nozzle diameter being fixed as that of the BR 
and ASR) and hence gave low gas velocities and 
poor low-temperature performance, a 1.75-lb charge 
was standardized as the Mk 16. 

Occasional difficulties in low-temperature static 
proof firing of the Mk 16 grain together with the 
fact already mentioned that the internal K was 
higher than desirable for good high-temperature 
performance made it advisable to design a new 
grain which would be slightly longer and slimmer. 
This would give a higher nozzle K and a lower in- 
ternal K, thus improving performance at both ends 
of the temperature scale (see Chapter 22) . Dimen- 
sions were changed from 1.70x0.28x12.5 in. to 
1.66x0.26x13.25 in., and the latter grain was 
standardized as the Mk 16 Mod 1. Although cellu- 
lose acetate end washers on the ASR and BR 


178 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


grains had been discarded as unnecessary, they are 
used on both ends of the SCAR grains, which other- 
wise would have been too regressive on account of 
their large web thickness. 

The Mk 17 grain for the Model 3 is simply a 
shorter version of the Mk 16 Mod 1. So much dif- 
ficulty with ignition was experienced with it at low 
temperatures, mainly because of the large empty 


those for CIT production were shaped cold in one 
piece from tubing. With these nozzles, no difference 
in dispersion from that of the machined nozzles 
could be detected. At the request of several Navy 
contractors, CIT ran numerous tests of nozzles 
made in two parts in punch presses and held to- 
gether by a press fit at the throat. This method of 
fabrication left a small step at the exit end of the 


NOZZLE CLOSURE INHIBITOR DISK 


PROPELLANT GRAIN 


FRONT CLOSURE DISK 


NOZZLE GRID 


SUSPENSION BUTTON 


MOTOR TUBE 


SPACER 


SUSPENSION BUTTON 



SHORTING CLIP ELECTRICAL CONNECTOR CABLE AND PLUG ALTERNATE NOZZLES 


Figure 7. Subcaliber aircraft rocket. 


space at the front of the motor, that the standard 
12-g BR igniter was superseded by one containing 
14 g of powder. 

Nozzles. The BR nozzle was used first because 
it was a standard item already in production, but, 
when the Navy production began, different con- 
tractors were used so that this was no particular 
advantage. Consequently most SCAR’s have been 
made with brazed-in formed nozzles because they 
are cheaper. The formed nozzle was patterned 
closely after those of the 3.25-in. AR motor, and 


nozzle throat, just in the critical position to affect 
the gas flow in the exit cone, and probably for this 
reason SCAR’s containing these nozzles had dis- 
persions twice that of CIT rounds or more. They 
were thus never approved for production, although 
it was repeatedly pointed out that the trouble could 
probably be cured by a simple change in design to 
put the step on the entrance side of the throat. 
Nozzles made in two half shells (i.e., split in a plane 
through the axis and brazed together) performed 
satisfactorily in their only field test. 


5.0-IN. HIGH-VELOCITY AIRCRAFT ROCKETS [HVAR] 


179 


Lugs. The lightness of the round made it unnec- 
essary to provide large lugs for sway bracing, and so 
a specially shaped lug button was designed (see 
Figure 7). It had a head which rode on the top of 
the Mk 4 launcher slot and a wide shoulder which 
fitted fairly closely below the slot. It was entirely 
satisfactory, and the only troubles were with the 
method of attaching to the tube. Methods tried 
were (1) threading them into the tube and silver- 
soldering, (2) arc-welding them into unthreaded 
holes in the tube, and (3) attaching a special flux- 
filled stud with a special welding gun and upsetting 
the end of the stud to hold the lug button on. The 
last method was by far the cheapest, quickest, and 
generally most satisfactory, and replaced other 
methods for CIT production as soon as it was tried. 

Fins. The radial fins are spot-welded together 
and the assembly spot-welded to the tube in a sim- 
ilar manner to the CWG. This is satisfactory since 
no significant saving of space would be made by 
having them detachable. 

Heads. The original Mk 1 head was machined 
from steel and weighed 1.6 lb. The shortage of 
steel led to a request from the Bureau of Ordnance 
to investigate the possibility of using die-cast zinc 
heads. Since the head is situated where the gas 
stream is essentially stagnant, it was found that 
zinc heads do stand up satisfactorily in general, but 
in at least two cases a little leakage through the 
threads occurred and the gas eroded a hole about 
1 in. in diameter before the end of burning. Al- 
though it was difficult to reproduce the phenomenon 
at will, it appeared that out-of-round motor tubes 
might cause it and that proper luting of the threads 
would prevent it. Accordingly it was specified that 
the head and motor threads be coated with a non- 
drying luting compound known as “Crater Com- 
pound. ” The zinc heads were designated Mk 1 
Mod 1 and Mk 3. The heavy Mk 2 head for the 
Model 2 SCAR w T as made only from steel, and its 
final weight was 8.6 lb. Several other weights were 
tried previously in attempting to get the trajectory 
correct. 

1933 Launchers 

The SCAR’s were designed to be fired from the 
Mk 4 rails without modification, but were too 
short to reach between the posts of the Mk 5. 
Various adapter launchers were tried having dif- 
ferent lengths from zero up to about 3 ft, since it 


was possible to control the tip-off and gravity drop 
by the launcher length and thus get the best fit to 
the trajectory of the standard rounds. A 2-in. 
constrained travel of the front lug was finally 
adopted, and this adapter was standardized as the 
Mk 6. 

1934 Reports 

Many of the reports on sight settings, trajectories, 
and use of other aircraft rockets contain information 
on the SCAR’s. The only formal reports on the 
rockets themselves are references 31 and 32. 

194 5.0-IN. HIGH-VELOCITY AIRCRAFT 
ROCKETS [HVAR] 

The 5.0-in. AR with the 3.25-in. motor was from 
the time of its inception admittedly a stopgap. Its 
velocity of only 700 fps gave it too little penetrat- 
ing power and too much gravity drop and required 
that, to be effective, it be fired at relatively short 
range where antiaircraft fire was uncomfortable. 
In addition, its lack of stability under water re- 
stricted its usefulness as a Navy weapon. To accel- 
erate the same 50-lb 5.0-in. head of the 5.0-in. AR 
to a velocity equal to or greater than that of the 
3.5-in. AR required obviously a motor of larger 
caliber. By the late summer of 1943, extrusion 
presses were available which could make consider- 
ably larger grains than 3-in. diameter, and shortly 
after the design of the Mk 13 cruciform grain had 
been stabilized the propellant section began experi- 
ments on possible grains for a 5.0-in. motor, inside 
diameter 4.625 in. 

As expected, the 5-in. grains gave the same 
answer as the 3-in. grains; namely, that for high 
loading density the cruciform shape is considerably 
superior to the tubular and that a spiral inhibiting 
pattern gives satisfactory burning curves. A 24-lb 
grain was designed, having a web thickness of 1.6 
in. and an outside diameter of 4.22 in., so that with 
0.15-in. thick inhibitor strips it was a reasonably 
snug fit in a 4.625-in. tube. This grain, designated 
Mk 18, gave beautiful neutral-burning pressure- 
time curves at all temperatures from — 25 F to 160 
F. Performance was so good that it was believed 
that a 20-per cent heavier grain would still work 
satisfactorily, but the difficulties with the 3.25-in. 
AR motor had taught us that it did not pay to try 


180 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


to push to the limit of grain size, and 24 lb was 
adopted as standard. This amount of propellant 
was more than enough to give a faster round than 
the 3.5-in. AR. 

The first 5.0-in. motor came off the drawing board 
in early December 1943, and probably underwent 
fewer significant design changes in the course of its 
development than any other rocket motor. The 
conservative design of the charge paid ample divi- 
dends. In field firing at 140 F, even in some cases 
with heads weighing only 20 lb which gave consider- 
ably larger accelerations than normal, malfunction- 
ing was so rare even with the ordinary JPN powder 
that this temperature was adopted for regular proof 
firing. Its low r -temperature performance was amaz- 
ing. Because of the blowout disk which enabled it 
to run at a A of 216 (for the 3.25-in. AR motor, 
K = 167), it practically has no low-temperature 
limit. To the author’s knowledge, it has never been 
known to chuff, and even the two rounds which 
were packed in “dry ice” at — 110 F over night 
showed no evidence that they were near the failure 
point. 

From the beginning it was nicknamed “Holy 
Moses,” obviously because at the time it appeared 
it seemed like such an enormous rocket. Since a 
number of apocryphal versions of the circumstances 
under w T hich it got its name are current, the 
author may be pardoned for setting the record 
straight . It is said, for example, that “Holy Moses” 
was the exclamation of the first pilot who fired one. 
The fact is that, before it was even off the drawing 
board, the author gave it that name as an experi- 
ment to see if he could make it stick and become 
the universal unofficial name. It did. 

The design and development of the Holy Moses 
motor (the 5.0-in. Mk 1) was completed about 
June 1, 1944, and CIT production was in the 
process of changing over from the older CIT Model 
1 motor when “Project Moses” appeared on the 
scene. The V— 1 “buzz-bombs” had just begun 
falling on England, and the fundamental strategy 
of resisting them was to eliminate the launching 
sites by aircraft attack, especially those in the Pas 
de Calais area. It was thought that the HVAR 
might prove an effective weapon against them, and 
it was suddenly decided to begin approximately five 
days later shipping the entire CIT production (100 
rounds per day) by air to England. Fifty sets of 
launchers were also included, and a special mission 
accompanied them to England to equip a squadron 


of P-47 fighter planes for service-testing of the 
equipment. Nineteen shipments of 100 rounds 
each were made by air, together with one boat ship- 
ment of 500 rounds of the obsolete experimental 
ammunition which could be scraped together. By 
the time the 513th Fighter Squadron (SE), 406th 
Fighter Group, Ninth Air Force, AAF, was equipped 
and ready for training, it had been determined that 
the launching sites were not suitable rocket targets, 
and the ammunition was available for supporting 
the invasion of France, which it did with excellent 
results. In a letter of commendation written to 
NDRC, Major General B. E. Meyers stated that 
this initial use of the HVAR proved “without ques- 
tion the effectiveness and efficiency of this equip- 
ment in actual combat, and has resulted in providing 
the Army Air Forces with the best antitank weapon 
of the war.” 

The combat experience in Normandy emphasized 
two facts that were already known: (1) that the 
post launchers designed for the smaller AR’s were 
not rugged enough for Holy Moses, and (2) that 
the inferior armor-piercing qualities of the head 
was a serious disadvantage. The AAF was suffi- 
ciently impressed, however, to adopt the rocket as 
standard fighter plane equipment and to undertake, 
in cooperation with CIT, a high-priority program 
of launcher development, so that by the spring of 
1945 some Army fighter planes began coming off 
the production lines equipped to fire the Navy’s 
HVAR. 

Two views of the assembled rocket are shown in 
Figure 5 of Chapter 17. 

Design Features 

Tubing. NE 8735 steel was specified for the 
motor tubing, and, as in the case of the 3.25-in. 
AR motor, it was specified by internal diameter 
(4.625 ±0.015 in.) and minimum wall thickness 
(0.187 in.). Since the tubing received averages 
thicker than 0.200, the motor is somewhat heavier 
than necessary, but this is a minor disadvantage. 
In other respects the tubing is almost ideal. To the 
author’s knowledge, no motors were rejected for 
failing the pressure test at 5,000 psi although CIT 
production exceeded 100,000, and no motor bursts 
occurred in field firing which appeared to be the 
fault of the tubing. Standard motor tubes burst at 
pressures between 6,000 and 7,000 psi. Field tests 


5.0-IN. HIGH-VELOCITY AIRCRAFT ROCKETS [HVAR] 


181 


with tubing more than 30 per cent stronger showed 
that increasing the tubing strength had no effect on 
the frequency of bursts at 160 F. Except for facing 
the ends and threading, the only machining on the 
tube is the counterbore at the front end to provide a 
close fit to the guiding land on the head and to the 
front motor seal. The first model had a longer tube 
and a correspondingly longer counterbore to accom- 
modate a different head, as explained later. Be- 


thinner grid. The eight peripheral nozzles and one 
central blowout nozzle are machined in the solid 
nozzle plate because the nozzle area is too large to 
permit insert nozzles. The tooling necessary to 
make such a nozzle plate with sufficient accuracy 
and to check it for alignment is rather complicated, 
and during the first three months the accuracy of 
the rocket steadily increased as nozzle production 
technique improved. 



cause of the known disadvantages of internal motor 
threads, particularly Y threads, pressure tests 
with square threads were made, but it was decided 
that their use would unduly complicate production 
and gauging. 

Nozzle. The design of the nozzle and associated 
parts can be seen in Figure 8, and a rear view of the 
nozzle is shown in Figure 5 of Chapter 17. It is 
based upon the experience with the 3.25-in. multi- 
nozzle motor but includes one entirely new feature 
in supporting the grid on a grid stool in the center 
of the motor, thus allowing the use of a much 


The nozzle ring or skirt extending beyond the 
rear face of the nozzle plate serves as a receptacle 
for the electric plug during shipment, but its 
primary purpose is to reduce the luminosity of the 
gas in the same way as a large expansion ratio 
does for single nozzle motors. 

The grid stool serves three purposes: (1) it sup- 
ports the grid, which is cemented to the propellant 
grain, (2) it clamps the blowout disk in place, and 
(3) it allows the motor pressure to get to the blow- 
out disk while holding the insulation to protect it 
from the heat and erosion of the gas. It was found 


182 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


that the effective gas velocity decreased consider- 
ably when the blowout disk functioned, and it was 
thought that a redesign of the grid stool to allow 
more direct access of the gas to the central nozzle 
might improve it. A cast steel stool, square in 
cross section, was designed which has four gas 
access holes inclined at an angle instead of being 
perpendicular to the rocket axis. No difference in 
gas velocity could be detected, and it is probable 
that the turbulence of the gas flowing through the 
central nozzle is only part of the reason for the 
reduced efficiency, the lower nozzle K and nozzle 
coefficient also contributing. A later design than 
that of Figure 8 has the blowout disk in the form of 
a shallow copper cup which is crimped onto the grid 
stool, thus making it impossible to insert two disks 
in the same motor. 

Suspension Lugs. Welded lugs were chosen in 
place of lug bands for a number of reasons. The 
spacing between launcher posts had been fixed by 
the 3.25-in. motor and did not appear likely to 
change; nor were American rockets being fired from 
British launchers, so that the arguments which led 
to the use of detachable lug bands on the 3.25-in. 
motor no longer held. Fixed lugs had the important 
advantage of rigidity and invariable spacing, and in 
addition they made possible an appreciable decrease 
in air drag. Since the Mk 4 launcher was by now 
obsolete, the lugs were made to fit post launchers 
(see Figures 3 and 4 of Chapter 17), although a 
small attachment was made to fit on the rear 
“tunnel” lug for use with the Mk 4. It was used 
very little, if at all, and was not even made in 
Navy production. Since even in ground firing, long 
launchers give no appreciable decrease in dispersion, 
there is little reason for their use. 

Fins. In order to fit the same launchers, the fins 
were made the same size as those on the 3.25-in. 
motor. Detachable fins were decided upon because 
the motors were so heavy that they had to be boxed 
individually, and a one-piece tail like that of the 
Mk 7 motor increases the shipping volume per 
motor by more than 35 per cent. The fins (see 
Figure 4 of Chapter 17) were therefore die-formed 
in two pieces and seam-welded together at the edges, 
leaving a hollow space % in. thick inside to house 
the latching mechanism. Details of the latching 
mechanism are shown in Figure 8. It was found to 
be quite satisfactory. The dimensions had to be 
worked out by trial and error, but when fins and fin 
lugs were properly made and not fouled with paint 


(a point which had to be watched), the fins were 
easy both to install and to remove and fitted very 
tightly. All the latch and lug parts could be 
stamped from sheet, so that they were not ex- 
pensive. The four fin lugs were welded to the 
motor after the nozzle was installed, and no dif- 
ficulties with this procedure were found. 

As previously pointed out, the choice of 5x8 in. 
for the fin dimensions was arbitrary, being simply 
copied from the aircraft version of the RP-3. In 
the summer of 1945, a comprehensive test of 
possible HVAR fin shapes was made in the high- 
speed water tunnel at CIT, 33 and among the results 
were the following: 

1. For 5-in. width, 8-in. length is very close to 
the optimum from the standpoint of accuracy. 
Ten-inch length would give very slightly more 
stability, but 15-in. length would be worse. In 
general, for any width, increasing the length be- 
yond about 1.5 calibers gives little or no increased 
stability. 

2. The stabilizing moment increases very rapidly 
with an increase in fin width. Thus 8 x 8-in. fins 
would reduce the yaw oscillation distance from 320 
to 240 ft and reduce the dispersion from a zero- 
length launcher in the same ratio. 

3. Tests with ring tails were made also, even 
though they cannot be used with post launchers. 
It was found that for a given size ring tails are much 
more efficient in providing stability than fin tails. 
This is illustrated in Figure 9 which shows six dif- 
ferent tail shapes, all of which would give the same 
yaw oscillation distance (and hence the same dis- 
persion, presumably) as the standard tail shown 
in the lower left. 

4. The stabilizing moment for a given tail is 
quite constant for overall rocket lengths between 
10 and 14 calibers, so that the results should be 
applicable to other shorter rockets of uniform 
diameter (such as Tiny Tim) . 

Igniters. Pending the development of a larger 
igniter, the earliest 5.0-in. motors contained two of 
the 35-g capacity plastic case igniters which were 
used in 3.25-in. motors. These gave satisfactory 
ignition but were not satisfactory for service use 
because they were not held securely in position. A 
rather heavy plastic case igniter 4.56 in. in diameter 
was tried. In order to accommodate two squibs 
and their connecting wires in a squib compartment 
at the rear of the igniter case, the threaded cover 
for the powder chamber was put at the front end. 


5.0-IN. HIGH-VELOCITY AIRCRAFT ROCKETS [HVAR] 


183 


This proved to be a fatal flaw in the design, for, 
when the powder ignited and bulged the case walls 
out until they contacted the motor tube, all the 
burning powder found itself confined between the 
front motor closure and the heavy plastic piston 
formed by the igniter case. The resulting force on 
the grain fractured it and motor tube bursts occurred 
at least 20 F below the temperature at which they 



RING TAILS 


->i 0.4 k- 




-H 

0.8 \ 

<— 



\< 1.6 H 








( 

< 


a 

( 


o 

0 

r 

0 

r 

r> A 



> 

f 



l 




FIN TAILS 

Figure 9. HVAR tail designs giving the same 
stability as the standard according to water tunnel 
tests. 


had previously been found with the same weight of 
black powder in two small igniter cases. 

A 55-g metal case igniter was then designed and 
became standard. A 70-g metal igniter was found 
to increase the high-temperature burst frequency 
slightly. The extra space which had been left for 
the much thicker plastic igniter was taken up by 
inserting a thin cylindrical steel spacer. Since 
length is not undesirable in fin-stabilized motors, 


shortening the motor to remove this waste space 
was neither contemplated nor tested. It was 
simply left as insurance against changes in propel- 
lant length. A certain amount of space is probably 
necessary to get good ignition, but the problem did 
not arise in the case of the 5.0-in. finner motor and 
so no tests were made of it. 

When the Tiny Tim igniter was reduced from 
1,200-g capacity to 230-g in order to reduce the 
blast effect, the question of reducing the charge in 
the Holy Moses igniter arose. Thirty-gram metal 
case igniters proved to give but little less blast, 
however, and were inferior to the standard 55-g 
igniter at low temperatures, so no change was 
made. 

Seals and Closures. The design of a metal front 
end motor seal, later used in the 5.0-in. spinners 
and shown in Figure 13 of Chapter 23, was first 
developed for the HVAR. For this motor the seal 
had a well in the center so that the blowout patch 
was recessed from the front face, leaving a space 
for the cap or bracket on the base fuze. Glued to 
the seal are a 1-in. felt on the back side and a 3^8-in. 
felt on the front side, both perforated so as not to 
interfere with the blowout patch. The seal is in- 
serted with a tool which positions it accurately 
so that the head, or the thread protector which 
extends into the motor the same distance as a head, 
seats against the thin felt washer and keeps the 
seal from shifting and breaking loose. A thin steel 
cup is sealed in the thread protector to provide an 
auxiliary seal at the front end. 

The rear auxiliary seal, which as in the case of the 
3.25-in. motor serves also as a receptacle for the 
pigtail, is pressed into the nozzle skirt ring. It was 
made dome-shaped in order to make it impossible 
to stand the motor on the nozzle end. 

Heads. The first head, which eventually became 
the Mk 5, was made from the same 5.0-in. A A shell 
which had given the 5.0-in. AR its head. The only 
changes made on it were to bore out the base to 
take the PIR fuze and to thread the outside to fit 
the motor. For extra support in oblique water and 
ground impacts, the head thread was moved for- 
ward so that the base of the head extended into the 
motor tube and carried a “guiding land” machined 
to fit closely (minimum clearance 0.010 in. on the 
diameter) into the counterbore in the tube, which 
had the same diameter as the minor diameter of the 
motor thread. Original experimental models had a 
5.5-in. overlap of the head and motor, but this 


184 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


was reduced to 3.0 in. because the longer overlap 
was believed to make the wall thickness of the head 
too small at one point. The final base design is 
shown in Figure 1 of Chapter 15 and was used on 
all HYAR heads. 

The Mk 5 head was used because of its easy 
availability, and it has a number of serious draw- 
backs. Because of its relatively thin wall and the 
large hole in the nose, it does not perform well 
against concrete or armor plate but breaks up 
easily. It is also unstable under water and under 
ground because of too great nose lift. Three other 




Figure 10. HVAR heads. Top: service head Mk 
5 or Mk 6 with Mk 149 nose fuze and Mk 157 
(PIR) base fuze (early design, prior to adoption 
of gas-check ring). Below : heads contemplated for 
service use. Modified Mk 46 shell (similar to Mk 
2 head) with windshield and Mk 166 (DDR) base 
fuze. Model 31 (similar to BuOrd Type Ex-1) 
with AIR-12 nose fuze and Mk 166 (DDR) base 
fuze. Model 32 with Mk 166 (DDR) base fuze 
only. 


heads were therefore designed and had been partially 
tested before the end of World War II. They are 
shown, with the standard Mk 5, in Figure 10. 
The CIT 5.0-in. Model 35, essentially the same 
as the BuOrd 5.0-in. Mk 2, was designed after the 
5.0-in. special common projectile Mk 38 or Mk 46 
for armor piercing. Preliminary tests indicated that 
it would penetrate 2-in. STS plate at up to 40 


degrees obliquity and would probably penetrate 
3-in. plate at normal incidence if the pyrotechnic 
delay in the base fuze were long enough. The CIT 
Model 31, similar to BuOrd Type Ex-1, was de- 
signed on the basis of water impact tests to have 
optimum underwater trajectory and was later 
found to have optimum underground performance 
as well. Although the water-discriminating fuzes 
originally intended for use with this head were 
abandoned, it would be much better for general 
use than the Mk 5 if it had a DDR base fuze and 
an instantaneous nose fuze with the same hemi- 
spherical shape as the AIR-12. In oblique im- 
pacts on fairly heavy armor plate, the nose fuze is 
broken off, so for some purposes the Model 31 
should be replaced by the Model 32 having the 
same shape but with a solid nose. Although no 
plate tests have been made with this head, it ap- 
pears likely to be very useful if made from the 
proper steel. 

Fuzes. The Mk 149 was the only nose fuze used 
on the HVAR in service. Proximity nose fuzes 
were found to be unsatisfactory because of the 
prolonged afterburning, probably caused by the 
inhibitor strips, which are too small to be ejected 
through the nozzle as in the case of the 3.25-in. 
motor. 

The nondelay base fuze Mk 146 was used first 
but was replaced by the Mk 157 with 0.02-second 
delay. When the gas check ring was adopted, the 
fuze became Mk 159 and a shorter delay (0.015 
second) was used because of the increased velocity 
of the HVAR over the 5.0-in. AR for which the 
Mk 157 was originally designed. The Mk 159 in 
turn gave way to the Mk 164 which incorporates an 
improved shutter design to decrease the probability 
of duds at high impact angles. The DDR fuze, 
which was put into production but did not reach 
service use, is designated Mk 166. Description of 
these fuzes is given in Rocket Fuzes , 34 

Alternative Designs 

Nonwelded Versions. As a result of difficulties 
with welding fin lugs on the Tiny Tim motor tube, 
a decree was laid down by someone in the Bureau 
of Ordnance that no welding was to be permitted 
either on the Tim or the Holy Moses motors. The 
5.0-in. motor was therefore hastily redesigned in 
Washington, and the 5.0-in. Motor Mk 2 Mod 3 
became the standard model for Navy production. 


5.0-IN. HIGH-VELOCITY AIRCRAFT ROCKETS [HVAR] 


185 


It has two lug bands similar to those on the 3.25-in. 
AR motor and a tail which is attached to the motor 
by clamping the cylindrical portion with nuts and 
bolts. The two-piece hollow design of the fins 
themselves was maintained so that the tail has 
adequate strength. Its chief difficulty, the amount 
of shipping space required, was not felt to be 
important. 

At the request of the Bureau of Ordnance, CIT 
designed and made preliminary tests on another 
nonwelded model, which was designated only by 
its drawing number, 5MA4. In this design the fin 
lugs, rear suspension lug, and nozzle skirt are made 
in one assembly and attached by drive screws into 
the nozzle plate. This design allows the use of the 
individual detachable fins of the Mk 1 motor. 
Aside from this, the only important change was to 
redesign the lug band clamping system so that the 
band can be tightened more securely and to sub- 
stitute flat-bottomed positioning holes and pins for 
the tapered ones which had been used on the 3.25- 
in. Mk 7 motor and carried over to the 5.0-in. Mk 3. 
These changes position the front lug band securely 
enough so that there is no danger of slippage under 
the stresses normally applied in service. No 5MA4’s 
were produced. 

CIT produced more than 100,000 Mk 1 motors 
without any difficulty with welding on the motor 
tubing. Failures occurred only at extremely high 
temperatures and always, as nearly as could be 
determined, as a result of grain failure. Occasionally 
such bursts showed a tendency to occur along one 
of the welds on the fin lugs because of the slight 
weakening at this point, but equally often the split 
ignored the welds entirely. 

White Whizzer. In response to a Navy request for 
an experimental 5.0-in. motor to give the highest 
possible velocity, the 5.0-in. Motor Model 38 was 
designed. It was nicknamed the “White Whizzer’ ’ 
after the author’s favorite football player, “Whizzer” 
White. The use of the motor was not originally 
specified, but it proved to be for the purpose of 
accelerating the ram jet motor which was being 
developed in the East at JAV-APL (Sec T) . It was 
not felt desirable to use a longer grain than the 
Mk 18 unless absolutely necessary, and so the Mk 1 
motor was simply lightened as much as possible. 
The motor tube was shortened by 5 in. and ma- 
chined on the outside (except at the ends) to a wall 
thickness of 0.125 in., thus reducing its weight from 
44.7 to 27.7 lb. The grid stool was lightened and 


shortened by eliminating the blowout disk, and 
some metal was removed from the nozzle plate to 
lighten it slightly. Suspension lugs were omitted 
and small lightweight fins, attached to a cylinder, 
were held in place by bolts into the nozzle plate. 
The result was a loaded motor which weighed 62.2 
lb instead of the standard 88.3 lb. With the stand- 
ard HVAR payload, its velocity was almost 50 per 
cent greater than the HVAR, and with light heads 
it was actually clocked at 2,490 fps. This velocity 
requires an acceleration in excess of 100 g, so that 
the force on the grain would certainly restrict the 
upper-temperature limit seriously, but no difficulty 
was found with it up to 100 F, which was the 
highest temperature at which it was tested. No 
information is available concerning the Navy’s 
use of the motors which were supplied by CIT. 

19,4,3 Launchers and Service Use 

The launchers for the HVAR are the same as for 
the AR’s except that its greater weight necessitated 
more rugged designs and impelled the change from 
Dural to high-tensile steel for post launchers, as 
mentioned in Chapter 17. 35,36 

After its first spectacular and successful test in 
Normandy, the HVAR was very little used by the 
Army because of failure to set up any adequate 
and comprehensive program of pilot training and 
failure to coordinate supply so that the rockets were 
available at the times and places where they might 
have been effectively employed. This situation was 
in the process of being remedied when World War II 
ended. With the Navy in the Pacific, the HVAR 
gradually supplanted the 5.0-in. AR as it became 
available. As anticipated, it proved to be a great 
improvement over the slower 5.0-in. AR, but the 
details of its use must be found in Navy publications. 

1944 Reports 

On the ammunition itself, the two most important 
CIT reports are references 37 and 38. Various 
aspects of its use in forward firing are discussed in 
many of the reports listed in Section 19.2.7. Manu- 
facturing problems are treated in references 39, 40, 
and 41. Motor-loading procedures, applicable 
essentially either to HVAR or “White Whizzer,” 
are detailed in reference 42. 


186 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


195 11.75-IN. AIRCRAFT ROCKETS 

The much better accuracy and penetrating power 
achievable with forward-fired rockets than with 
bombs made desirable the development of an air- 
craft rocket which could carry a payload compa- 
rable to that of a large aircraft bomb. Sporadic tests 
of accelerating standard bombs with several small 
rocket motors had been made from time to time at 
CIT and elsewhere, but this was a clumsy and 
inefficient method of getting velocity and proved 
also to be very inaccurate. The obvious solution 
was one big rocket motor. Such a big motor became 
possible as soon as the 4.2-in. cruciform grain was 
available, and the development of the 11.75-in. 
aircraft rocket began in March 1944, soon after 
that of the HVAR. For obvious reasons, it was 
immediately nicknamed “Tiny Tim. ,, The first 
field firing was made on April 26; one static firing 
of the propellant charge had been made two weeks 
earlier. The design was logically developed from 
the 5.0-in. HVAR and presented a number of prob- 
lems not previously encountered in the project’s 
work with smaller rockets. These included: 

1. The use of a multiple-grain charge, which 
necessitated an internal structure for its support. 
Four Mk 19 cruciform grams, 60 in. long, were 
used, giving a propellant weight greater than total 
weight of a loaded HVAR. 

2. The use of threads on the motor much larger 
than, and different in shape from, those in standard 
commercial use which can be made in ordinary 
machine shops with commonly available tools. 

3 . The requirement of special devices for handling 
and attaching these larger rockets to airplanes. 

4. The large blast effect, which required (a) care- 
ful engineering to minimize, (b) special launching 
devices to separate the rocket from the airplane 
before ignition, and (c) a considerable program of 
research into the sighting and aiming problems of 
this type of launching. 

The Navy’s 500-lb SAP bomb AN-M58A1 ap- 
peared to be the most desirable head for such a 
rocket, and fortunately there was a standard oil well 
casing of the same diameter, 11.75 in. OD, which 
had adequate wall thickness and tensile strength 
and enabled the development program to get 
started without waiting for a special mill run of 
tubing. There was not much of it available, how- 
ever, and we were reduced for a time to the expe- 
dient of salvaging it from abandoned oil wells. 


Because of its size, which made production slow 
and posed extraordinary difficulties both in motor 
design and in installation on aircraft, the Tiny Tim 
was a long-term project in comparison to its pred- 
ecessors. Nevertheless, its progress was very en- 
couraging, and, when in June successful air firings 
began, it was decided that Tim was a likely supple- 
ment for the Holy Moses against the robot bomb 
launching sites. Thus on June 28, 1944, six days 
after the first air firing of Tiny Tim, a memorandum 
from the Navy Chief of Staff to the Vice Chief of 
Naval Operations assigned top priority to the devel- 
opment of the rocket and its associated launchers 
for the purpose of getting it into service as soon as 
possible. Work was to start immediately on proto- 
typing launcher installations for the F4U and F6F 
aircraft, and the SB2C was later added to the list. 
Although the design of the internal motor compo- 
nents was not entirely settled, CIT undertook pro- 
duction of sufficient motors to be able to supply 10 
per day to the Services. Several hectic weeks fol- 
lowed before the high priority was deferred on 
August 7 because it became clear that the bomb- 
launching sites would be captured before Tim could 
be put into action. Two weeks later the crash of an 
SB2C in an experimental test caused a sudden halt 
and a complete re-examination of the program, and 
in the ensuing months the difficulties with blast and 
the various internal ballistics problems were studied 
in detail and gradually worked out. Development 
was essentially complete by October 1, Navy con- 
tractors began setting up for production, and the 
rocket was ready for combat test. Minor design 
changes and refinements continued for several 
months thereafter, however, dictated for the most 
part by the requirements of fitting to various types 
of aircraft. 

The following spring, aircraft squadrons with 
drop launchers were sent to the Pacific on the 
carriers Franklin and Intrepid for the first service 
test of Tiny Tim. The disastrous attack on the 
Franklin took place before its rocket planes ever 
went into action against the enemy, and the 500-lb 
explosive rocket heads in her hold contributed to her 
downfall. Although it is believed that the Intrepid's 
planes fired a few Tims against the Japanese, the 
Navy has not divulged any details. The Division 3 
history says they were used on Okinawa. 

Army Air Forces also undertook a program of out- 
fitting appropriate planes for firing the 11.75-in. 
AR. This program would have had the planes 


11.75-IN. AIRCRAFT ROCKETS 


187 


ready for action in the final invasion of the Japanese 
homeland. The end of World War II left Tiny Tim 
as a potentially powerful and effective weapon, 
which would enable a plane to deliver the punch of 
a 12.0-in. gun, but a weapon which never had a 
combat test of its capabilities. 

External Design Features 

The original specifications called for a rocket to 
be fired from aircraft having a 500-lb payload and 
as high a velocity as possible (preferably at least 

1.000 fps), and using as propellant four 4.2 x 1.5- 
in. cruciform ballistite charges. The rocket was 
to have multiple nozzles, including a blowout nozzle 
to increase its working temperature range, and for 
handling purposes it was to be capable of standing 
on its nozzle end. The first guess proved to be a 
good one on the two major components — the motor 
tube and the nozzle. Almost from the beginning 
their design was so stable that it was possible to 
continue regular production of them without con- 
sideration for the frequent and drastic revisions of 
internal design which occurred in the summer 
of 1944. 

Motor Tube. The choice of propellant fixed the 
internal diameter of the motor tube as not less than, 
and preferably not much more than, in. Its 
wall thickness was determined by the specification 
that it stand a 4,800-psi internal pressure test with- 
out permanent yield. Since saving weight was a 
primary concern, it was desirable to use tubing of a 
relatively high yield point in order to keep the wall 
as thin as possible. The grade N-80 API oil well 
casing, with an external diameter of 11.75 in., a 
0.489-in. wall thickness, and a minimum yield of 

80.000 psi, was the most suitable material found; it 
had the additional advantage of having the same 
outside diameter as the 500-lb SAP bomb which 
was being considered as a possible high-explosive 
head for the rocket. 

To obtain the required internal diameter it was 
necessary to machine the inside full length, and it 
was decided to machine the outside also, partly to 
save weight but primarily to assure accuracy. The 
10 per cent permissible variation in wall thickness 
could displace the center of gravity of the motor 
tube from the geometrical center of the ID by as 
much as 0.3 in., but it was desirable to keep the 
overall mechanical malalignment of the rocket as 


small as the gas malalignment, which with multiple 
nozzles was expected to be less than 1 mil (0.06 in.) . 

The diameters chosen, 11.7 in. and 10.9 in., with 
a minimum wall thickness of 0.380 in., give a 
maximum fiber stress of 76,800 psi (calculated by 
Barlow’s formula) for an internal pressure of 5,000 
psi. It was realized that this wall thickness was 
probably ultraconservative, since it was based 
upon standards evolved by the project from ex- 
perience with small motors which did not have a 
blowout disk to limit the maximum pressure in the 
motor. The fact that a burst of such a large motor 
would, it was believed, almost certainly result in 
destruction of the aircraft justified such conserva- 
tism, at least in the beginning. Later, tubing with a 
minimum yield of 90,000 psi became available and 
was specified by the Bureau of Ordnance for its 
production. Two high -temperature firings of Navy 
production motor tubes with walls reduced to 0.280 
in. were successful, and for the final production 
design (the Model 5 motor) a nominal wall thick- 
ness of 0.340 in. was specified. From the per- 
formance standpoint, considerably more drastic 
reductions could be made, as was further shown by 
later tests at NOTS, Inyokern, of motors with 0.200- 
in. walls. The increased velocity which can be 
gained by reduction below 0.340 in. is not very 
great, however, and for combat use from aircraft 
it is believed that a thinner wall is not desirable in 
view of its increased vulnerability to gunfire. 

The two ends of the tube were threaded inter- 
nally, one to take the body and the other to take 
the nozzle. In order to get as much strength at the 
threads as possible, the outside machine cut was 
stopped about 3 in. short of the ends. The thread, 
a modified buttress with a 3-degree loaded face, a 
50-degree included angle, and a pitch of 2}^, was 
designed for maximum strength against internal 
pressures combined with ease of assembly. The 
choice of 3 degrees was rather arbitrary; it was 
desired to keep the angle small in order to minimize 
the tendency of the end thrust on the nozzle to 
expand the motor threads, and 3 degrees was one 
of the common standard angles for buttress threads. 
When the prime contractors for large-scale produc- 
tion began making inquiries about the design, it 
became evident that the choice had not been the 
best one, since the smaller the angle, the smaller the 
diameter of a thread grinding wheel or hob which 
can cut the thread. From this point of view, an 
angle of about 7 degrees would have been preferable, 


188 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


since it allows the use of tools 3 to 3.5 in. in diameter 
and is still less than the angle of repose for friction 
between slightly lubricated steel surfaces so that 
there would be no tendency to expand the tube. By 
this time, however, production of bodies with this 
thread was already under way by the Naval Gun 
Factory, and the Bureau of Ordnance has not con- 
sidered the change desirable. 

A glance at an early general arrangement drawing 
of the 11.75-in. aircraft rocket reveals that there 
was 6 in. of empty space at the head end of the Mod 
0 motor. This came about as a result of a variety of 
factors. The original design of the structural mem- 
bers holding the propellant grains was such that it 
was expected that a considerable fraction of the 
gases would move forward and reverse their direc- 
tion at the front end of the motor, and adequate 
space was necessary to allow this to occur without 
excessive heating. The original head design had a 
12-in. overlap of the motor tube over the head for 
extra strength against oblique water impacts; the 
closure at the base of the body was a dome, convex 
forward, in order to leave the required space and 
still have the guiding land on the body as far back 
as possible. When the 11.75-in. Head Mk 1 was 
designed by the Bureau of Ordnance, the 12-in. 
overlap was reduced to 6 in. Because the design 
of the internal parts of the motor was so uncertain, 
it was decided not to reduce the length of the motor 
tube correspondingly. 

Subsequently, two factors appeared which made 
a reduction in length desirable: the interference 
between the tail of the rocket and the wing flaps on 
certain aircraft and the fact that the bomb elevators 
in a considerable number of aircraft carriers would 
not accommodate motors longer than 80 in., but 
required the use of other elevators for transferring 
the rockets from the magazines to the planes. At a 
conference in December 1944 with representatives 
of the Bureau of Ordnance, it was decided that the 
motor tube should immediately be shortened as 
much as possible without changing the service 
heads (Mk 1 and Mk 2) in order that the outside 
length of the motor shipping box could be kept 
under 80 in. Modification numbers were assigned 
for the shortened motor, and production of the Mod 
2 was begun by CIT as soon as new tubes could be 
made. It was subsequently found that the buttress 
thread was strong enough to stand water impact 
even without any overlap of the motor tube over 
the head. Consequently, in the final design 


(Model 5) the “skirt” on the head was removed and 
the motor tube was made as short as possible. 

As has already been pointed out, the Model 5 
motor was later found to be too weak to withstand 
ground impact and cannot be given a long under- 
ground trajectory even with the sphere-ogive head. 
Whether this is caused by the lack of overlap of the 
motor tube over the head, by the much thinner 
wall, or by a combination of the two factors has not 
been determined. Whether the Model 5 motor is 
actually an improvement on the previous designs, 
then, depends upon the tactical use. For most pur- 
poses, its higher velocity recommends it. 

Nozzle Plate. In view of the success of the 
HVAR, the choice of a multiple nozzle (see Figure 
11) for Tim was obvious. Various numbers of 



Figure 11. Nozzle end of Tiny Tim. 


nozzles ranging from G to 80 were considered. The 
choice fell upon 24 as a good compromise between a 
large number of small holes, which could be made 
with a simpler tooling, and a small number of large 
holes, which would be cheaper. The holes were 
arranged 1G in the outer row, 8 in the inner row, and 
one large one in the center, which is closed by a 
heat-insulated copper disk unless the pressure dur- 
ing firing rises high enough to eject it and thus 
increase the nozzle port area. Since the motor was 
to be capable of standing on end, the usual method 
of bringing the electrical lead out of one of the 
nozzles could not be used. Special electrical recep- 
tacles placed in the nozzle plate were therefore 
designed. A special electrical cable is supplied with 
each motor to go between the receptacle on the air- 
craft and those in the nozzle plate. 


11.75-IN. AIRCRAFT ROCKETS 


189 


Tails. Considerable evolution took place on the 
tails, but nothing need be said about the early 
designs because they were all based on the idea of 
welding lugs onto the motor tube. Unfortunately, 
the author, who was responsible for the design 
details, was not a good metallurgist, and it proved 
to be absolutely impossible to weld even so much 
as a %;-in. stud on the N-80 motor tubing without 
getting occasional failures in the pressure test of 
the tubing or in field firing. 

The tail which became standard was not of very 
elegant design and was never intended to become 
permanent, but Navy production began with it 
and World War II ended before the later improved 
design could be put into production. The individual 
fin pieces were made from J^-in. aluminum sheet, 
24ST, with radial beads rolled into the metal }/$ in. 
high for stiffening. The two halves were riveted 
together and to two steel bands which clamped on 
the motor tube, the rear band seating back against 
the ridge at the rear of the motor tube. The choice 
of aluminum over steel was made partly from 
weight considerations but chiefly because it was 
thought that less damage would be done to the 
propeller should a fin by any mishap get into its 
arc. The early fins were 12 in. wide, but in order to 
fit into the TBF bomb bay it w T as necessary to 
reduce them to 10x24 in., which became the 
standard. Interference with the wing flaps, which 
occurred with the adoption of the drop launcher in 
January 1945, caused the rear corner of the fins to 
be cut off, but, even with the corner removed, it was 
necessary on the Mod 0 motor to move the tail 
forward from its normal position in order to clear 
the flaps on some aircraft. 

For the Model 5 motor an entirely new tail was 
designed. It was considerably lighter than the 
standard and had individually attachable fins so 
that they could be shipped in the motor box with a 
consequent saving of about 10 cu ft of storage space 
per motor. Since the Model 5 motor was not pro- 
duced by the Navy, very few of the new tails were 
made, and still better designs have since been 
worked out at Inyokern. 

With regard to fin shape, the conclusions of the 
water tunnel tests on the HYAR are probably 
equally valid for the 11.75-in. AR, and consider- 
ably wider fins would be desirable if they would fit 
on the aircraft. Tests of telescopic fins have been 
tried at NOTS, Inyokern, and such fins might sig- 
nificantly increase the accuracy. 


Lug Bands. Lugs for ( attaching the motor to the 
airplane were originally ^welded to the motor tube, 
but this scheme had to be abandoned along with 
the welding of the fins, and the lugs also were 
attached to bands. This arrangement proved to be 
necessary for another reason, however, for it is 
impossible to use the same lug position on all air- 
craft. The Mod 0 motor was issued with the lug 
bands placed as required for the displacement 
launcher on the F4U, which was to have been the 
first installation to get into combat. Five bands 
were required: a standard bomb-hoisting lug at the 
center of gravity of the loaded round, two standard 
bomb suspension lugs to fit the standard bomb 
racks, and two launching lugs to attach to the dis- 
placement launcher and release the rocket at the 
bottom of the swing. In the latter part of 1944, 
tests on the drop launcher were so successful that 
the displacement launcher was declared obsolete 
and was removed from the airplanes. The drop 
launcher required only the three standard bomb 
lugs, but a second hoisting lug was attached at the 
center of gravity of the loaded motor for handling 
it before the head was attached. The change in lug 
band arrangement was made almost simultaneously 
with the change in motor tube length, so that almost 
all the Mod 0 motors had 5 bands, while almost all 
the Mod 2 motors and all Model 5 motors had 4 
bands. 

None of the lug bands made by CIT would stand 
up under the loads specified by the Bureau of Aero- 
nautics, corresponding to accelerations of IS Ag 
vertically (i.e., radially) and 11 Ag fore-and-aft. 
They were adequately strong for ordinary use, how- 
ever, and until the internal ballistics problems were 
resolved, there was no time to worry about lug 
bands. When comprehensive tests were made, it 
became apparent that it would not be possible to 
make suspension bands out of ordinary cold-rolled 
steel that would be strong enough to prevent slip- 
ping or distortion under the specified loads without 
a considerable increase in thickness over the % in* 
that had been used. The bands being made by CIT 
would take, on the average, only about half the 
specified loads, and those from the Navy contractor 
would take even less. Even the bomb suspension 
lugs themselves were too weak. Consequently, it 
appeared desirable to adopt heat-treated 4130 steel 
for the whole assembly and thus obtain parts about 
which no question of strength would exist. A 
minimum yield point of 100,000 psi was specified, 


190 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


and as this was two and one-half times the average 
of the cold-rolled %-in. bands, it was possible to 
reduce the thickness to }/$ in. and still increase the 
strength well above that required. Tests showed 
that J/g-in. suspension bands properly heat-treated 
would stand the vertical load test with a consider- 
able margin of safety and could be tightened on the 
tube so securely (75 ft-lb torque on 3^-in. bolts) 
that either band alone would withstand the specified 
vertical and fore-and-aft loads, although in actual 
practice the loads would always be divided almost 
equally between the two lugs. These bands were 
recommended for Bureau production. 

Internal Design Features 

Blowout Disk. The central nozzle is closed by a 
shallow copper cup, clamped in place by a threaded 
retainer. The cup (usually called a disk) is insulated 
from the motor gases by a 34-i n - asbestos-filled 
bakelite plug and a layer of hard-setting Permatex. 
Originally the disk was 0.064 in. thick and sheared 
at a hydraulic pressure (cold) of 3,000 psi. It was 
found that this disk did not always blow out at 
130 F, and, when it did not, high pressure peaks 
and much lower gas velocities were obtained in field 
firing. A reduction to 0.050-in. thickness, giving 
2,250 psi as the cold shearing pressure, raised the 
average gas velocity at 130 F from 5,430 to 6,340 
fps. Static-firing tests gave 3,120 ± 150 and 
2,490 ± 115 as the actual mean blowout pressures 
of the two thicknesses of disks, slightly higher than, 
but in reasonable agreement with, the values ob- 
tained with cold water pressure. The later adoption 
of JPN in place of JP propellant , b with the conse- 
quently lower pressure at high temperature, brought 
a further reduction of the disk thickness to 0.043 
in. in order to keep the safety factor of the motor as 
high as possible. 

Grid. The grid design was fairly obvious and has 
caused no difficulty except that it was originally de- 
signed much heavier than proved to be necessary. 
In trimming down the Model 5 motor to the mini- 
mum in weight, about 10 lb was saved by support- 

b The original ballistite composition used by CIT (standard 
trench mortar propellant was designated JP for “jet pro- 
pulsion.” In 1944 a slightly different composition became 
standard and was designated JPN (N for “new”). An experi- 
mental composition designed to have higher strength was 
called JPH (H for “hard”). All compositions contain roughly 
V 2 nitrocellulose and 3 /7 nitroglycerin with small amounts of 
other compounds. 


ing the grid on four legs instead of a ring. The ring 
was originally used to prevent erosion of the motor 
tube at the front face of the nozzle plate where the 
gases are deflected to go through the holes. This 
erosion had been found to be serious in the HVAR at 
high temperatures, but on the 11.75 in. it proved to 
be very small because of the difference in gas flow 
through the larger number of nozzles. To make 
doubly certain, the length of the motor tube 
threads at the nozzle end was made less in order to 
expose a minimum number of threads to the gases in 
front of the nozzle plate. That this change now 
made the two ends of the tube different was not 
thought to be a serious objection in large-scale 
production. 

Structure for Mounting Propellant Charge. When 
the idea of using a multiple-grain charge was 
advanced, enough experience had been gained on 
smaller grains, particularly the 2.74-in. cruciform, 
to indicate that they would not burn stably and 
smoothly unless each grain was shielded from the 
radiation given off by the others c and fairly well 
supported mechanically along its whole length. It 
was also desirable that the grains be held firmly 
down against the grid even under backward accel- 
erations of 12 g. The most persistent and difficult 
design problems arose in connection with the struc- 
ture for accomplishing these ends. 

Charge Support. Although, strictly speaking, 
nearly every internal part is a support for the 
charge, the name has been given to the structure 
which attaches to the grid at the rear end and 
serves to hide the grains from each other, supports 
them along their length, and attaches at the front 
end to the clamp which prevents the grains from 
moving forward. Only major variations in the 
charge support will be discussed here, since small 
changes were almost innumerable. 

The earliest tests, with charge supports which 
completely surrounded each grain, were unsuccess- 
ful because such supports had to be made out of 
fairly thin steel (0.075 in. was used) in order to fit 
into the tube. Flight of the rocket was satisfactory, 
but virtually the entire rear end of the charge sup- 
port was eroded away by the time the burning was 
three-quarters complete so that the grain broke up 
early and gave low gas velocity. 

c Recent research at Inyokern has shown that radiation 
effects are actually not serious in this case, however, so that 
considerably lighter and simpler designs of charge support 
can be made. See Figure 12C. 


11.75-IN. AIRCRAFT ROCKETS 


191 


The first successful charge support was the so- 
called “X type” shown in Figure 12A. It was 
formed from %-in. steel sheet and welded to the 
grid. It was realized that, touching the grain only at 
the corners, it might not give sufficient support, 
but it was simpler to make than other types which 
had been suggested, and the initial experimental 
tests were successful. It is probable that powder 
having a compressive strength as high as that of 
JPH would perform about as well in this charge 
support as in any other. When, however, a large 
quantity of ballistite was received with too high a 
nitroglycerin content and a consequently lower 
compressive strength, trouble was immediately en- 
countered in high-temperature proof firing. The 
new powder gave high pressure peaks and excessive 
powder breakup, and on one round an effective gas 
velocity of only 4,620 fps was obtained. In the 
belief that the difficulty was probably insufficient 
mechanical support of the grain, tests were begun 
with a new charge support, the 4Y type shown in 
Figure 12B. 

The success of the 4Y type in eliminating the bad 
high-temperature performance with 44 per cent 
nitroglycerin JPN propellant was spectacular. In 
one field test, it increased the gas velocity at 130 F 
by more than 1,200 fps and completely eliminated 
the end breakup peak as far as could be ascertained 
from the photographic data. The dimensional toler- 
ances as originally laid down would have given the 
grains the same amount of support that they have 
in the 5.0-in. motor, in which the ends of all four 
arms are supported and the spacing between sup- 
ports on opposite arms is very closely 4.625 in. 
This is accomplished in one direction by holding the 
arms of the Y’s accurately and in the other direction 
by holding the size and concentricity of the central 
square section so that the spacing between it and 
the ID of the motor tube is correct. It was never 
possible to meet these close tolerances in the fabrica- 
tion of the charge support, and the drawing toler- 
ances were progressively loosened to be in accord 
with the facts. In ordinary service, apparently, a 
very loose fit of the grain in the support is adequate. 
With powder of low quality or in high-temperature 
firings, one would expect the gas velocity and the 
number of failures to depend on the snugness of the 
fit. Therefore, the author has always taken the 
attitude (in discussions with Navy contractors) that 
it is worth a little extra trouble and expense to 
make the charge support as accurate as possible, 





Figure 12. Charge supports for 11.75-in. motor. 
Top : X type. Middle : 4Y type now standard. Bot- 
tom: Tubular type which may supplant 4Y type. 




192 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


even though it is impossible to prove experimentally 
that to do so will improve performance. 

The 4Y was welded to the grid and had four 
threaded studs at the front for holding the charge 
clamp. It was made from 11 -gauge steel in four 
sections which were originally spot-welded together , 
but, after some came apart during firing, a riveted 
design was tried and found successful. Elimination 
of the large amount of welding made it much easier 
to keep the parts straight and true. 

For Navy contractors who preferred them, how- 
ever, both spot-welded and arc-welded designs were 
included in Navy drawings as alternates. Also 
permitted was a design in which the charge support 
was bolted rather than welded to the grid, permit- 
ting the use of shims to get it concentric and coaxial 
with the grid. The arms of the Y’s at the nozzle 
end showed a tendency to warp away from the 
grains during firing, and they were reinforced by 
pieces of 1-in. angle iron. The only other change 
was to shorten the support at the time of moving 
the igniters from the rear to the front of the charge 
clamp which is discussed later. 

In the Model 5 motor an attempt was made 
to dispense with as much weight as possible. Vari- 
ous schemes were tried to make 4Y charge supports 
from 14- or 12-gauge material without success. It 
appears that nothing thinner than the standard 11- 
gauge material will hold its shape during firing well 
enough to give the grain its necessary support. 

Charge Clamp. The charge clamp is bolted in 
place as one of the last operations in motor loading 
to hold the grains firmly against the grid. In order 
to accommodate small length changes in the grains, 
about lj^-in. thickness of cruciform felt washers is 
placed between the clamp and the grains. Originally 
the igniters were placed between the felt and the 
front end of the grain, but, with the igniter size then 
in use, this arrangement was found to subject the 
grains and the clamp to rather large forces upon 
ignition, and the igniters were then mounted on the 
front face of the charge clamp. Various less rugged 
designs of charge clamp were tried and found to be 
too weak to withstand the ignition forces. The 
final design was a %-in. thick steel plate, torch-cut 
into the approximate shape of the four cruciform 
grains and bolted to four studs welded to the charge 
support (see Figure 13). With the igniter in front 
of the charge clamp instead of between it and the 
grains, the clamp can probably be made thinner 
and lighter if it should appear desirable. 


Igniters. The early experimental motors used 
either 16 or 24 Mk 9 igniters, which were developed 
for the 3.25-in. motor and contained 35 g of black 
powder each. As soon as it was available, the 
plastic case igniter which was developed for the 
5.0-in. Motor Mk 1 was adopted. This is 4.6 in. 
in diameter and 2.1 in. high, and has a powder 
compartment holding up to 200 g of powder and a 
wiring compartment for connecting the two electric 



Figure 13. Front end of charge for 11.75-in. mo- 
tor showing charge clamp and 230-g igniter. Later 
design eliminated exposed igniter wires. 


squibs. Four of these igniters were used, one in 
front of each grain, giving a total of 800 g of black 
powder. 

In the early firings, in order to get the grains to 
temperature it was necessary to remove the nozzle 
charge support assembly from the motor. As a 
result, in low-temperature firings, considerable frost 
formed on the surface of the powder grains. One 
static test showed an igniter peak of only 400 psi 
and a 50-millisecond delay in reaching the equilib- 
rium pressure of 850. Since in aircraft rockets it is 
desirable to have the pressure rise as rapidly as 
possible, it had been the general policy in deter- 
mining the adequacy of an igniter to have a pressure 
peak at the low-temperature limit about as high as 
the equilibrium pressure. On the basis of the — 35 F 
static performance it was decided to try a total of 
1,200 g of igniter, by filling the wiring compartment 
as well as the powder compartment and drilling 
holes between them. In the initial static tests the 
increased igniter appeared satisfactory from — 50 to 
144 F, and it was adopted as standard. Later the 


11.75-IN. AIRCRAFT ROCKETS 


193 


four plastic igniters were superseded by four tin 
plate igniters of the same capacity, also containing 
two squibs. From the beginning it was intended 
that the tin plate igniters should be used when 
available, but the abandonment of plastic igniters 
was accelerated by the discovery of a piece of plastic 
which had been blown into the oil cooler of an F4TJ- 
1D, incapacitating the aircraft. The new igniters 
were made of 0.010-in. tin plate on ordinary tin can 
machinery with top and bottom crimped to the 
sides with the standard “double crimp.” 

In experimental firings from wing launchers on 
the SB2C airplane, either with fixed or “lanyard 
drop” launchers, there was severe damage to the 
elevators. Investigation with high-speed cinema- 
tography disclosed that the elevators were given a 
severe and brief acceleration, piesumably by a 
shock wave, before the main propellant blast was 
set up . It was soon found that the magnitude of this 
shock wave is roughly proportional to the size of the 
igniter. Accordingly, it was decided that the igniter 
should be as small as possible even at the sacrifice 
of low-temperature performance, and a single tin 
can containing 230 g of black powder was adopted 
as standard. 

In retrospect, it is clear that, if the grains had not 
been frost-covered on the early cold shots, we would 
not have concluded that 1 ,200 g of igniter was neces- 
sary. The proper amount from the standpoint of 
good ignition is probably 800 g or somewhat less. 
When this factor is balanced against the shock 
wave damage to the aircraft, it is very difficult to 
determine the optimum amount to use. Tests con- 
ducted at NOTS in March 1945 on the effect of 
igniter size on blast damage showed that the main 
blast was larger than the igniter blast up to about 
500 g of igniter. It was therefore recommended that 
the igniter charge be doubled in the interest of 
better ignition at low temperatures. No such ig- 
niters had been made by the time the rocket was 
turned over to NOTS. As an alternative, two of 
the smaller igniters could be used, but this seemed 
undesirable since it increased the power require- 
ments and complicated the design. 

Igniter Leads. The method of connection and 
protection of the wires running from one or more of 
the igniters to the electrical receptacles in the 
nozzle plate was a persistent problem. Various 
troubles involved in making connection to four 
igniters will not be discussed. When the single 
igniter was introduced, the wires, which had for- 


merly been brought out near the outside of the tin 
can, were moved to the center and a 1.0-in. hole 
was bored in the center of the charge clamp to admit 
them into the central square in the 4Y charge sup- 
port. At the grid, the wires passed out of the 
central square through two rubber grommets (later 
combined into a single two-legged grommet) and 
thence to the receptacles. This arrangement was 
satisfactory except that the wires (about 10 ft of No. 
16 stranded copper, insulated) were always ejected 
during burning. In an attempt to keep them inside, 
a number of schemes were tried: wrapping the wires 
around the grid, tying them to a rivet at the front 
end, running them through small holes in a bulk- 
head at the front end of the square, and plugging 
the central square with a plastic material which was 
cast around the wires. The design of the Mk 1 
motor was frozen with no method of imprisoning 
the igniter wires. In the Model 5 motor, the wires 
were brought through the grid through small, snug- 
fitting holes without grommets. With this arrange- 
ment, almost all of the igniter leads remained in 
the motor during firing. 

Occasional motors were found to be short-cir- 
cuited because small flakes of steel and beads of weld 
dropped from the cracks in the charge support into 
the receptacle holes in the nozzle. To prevent this, 
the holes were filled with a plastic material. Several 
were tried, the best being “3-M Weather Strip 
Cement” (Minnesota Mining and Manufacturing 
Co.). 

The electrical leads from the nozzle plate to the 
aircraft also caused considerable trouble, particu- 
larly in drop launching, because the wind force 
tended to break them and because they had to be 
coiled so as not to tangle. In the final design (shown 
in Figure 11) the joint between the two-conductor 
cable and the two single-conductor cables, at which 
breakage usually occurred, was eliminated by un- 
raveling about 1 ft of the two-conductor cable, tying 
the individual insulated conductors in an electrician’s 
knot, stretching them into the form of a T, and 
molding rubber over them. Numerous schemes for 
coiling the lead were tested and rejected. The 
method finally adopted was to lay the excess cable 
along the motor tube in one long loop and attach 
the loop by means of special aluminum clips to the 
length of cable running from the nozzle plate to the 
suspension lug at the center of gravity of the round. 
This design materially reduced the number of mis- 
fires in drop launching. 


194 


DESIGNS OF FIN-STABILIZED ROCKETS FOR AIRCRAFT 


Motor Seals. The design of primary and auxiliary 
seals for the motor was relatively straightforward, 
based on experience with the 5.0-in. HVAR. Each 
nozzle is sealed individually with a die-formed steel 
cup 0.010 in. thick. The seal for shipping purposes 
on the Model 5 motor is a shallow steel pan screwed 
against a rubber gasket on the rear face of the 
nozzle plate. An attempt was made to reduce the 
mass of the 24 little sealing cups, but it appears 
that thinner cups do not give a reliable seal (0.005 
in. being entirely too fragile) and aluminum cups 
are destroyed in a short time by electrochemical 
action in a salty atmosphere. The front seal is a 


was forged in one piece. It was closed at the base 
by a steel plug accommodating three PIR or DDR 
base fuzes, and both the plug and the fuzes were 
sealed in place with gas check rings. It became the 
standard production head. For the Model 5 motor, 
the long “skirt” was removed from the Mk 2 head, 
making it the Mk 4. A sphere-ogive head (see 
Figure 13 of Chapter 24) was also designed and 
tested both under water and under ground. 
Although it had a much longer underwater and 
underground trajectory than the Mk 1 or Mk 2 
heads, the latter were also stable, and so the sphere- 
ogive head was not put into production. Most CIT 



Figure 14. 11.75-in. rocket ready for loading on drop launcher. 


very tight-fitting steel dome inserted with a hy- 
draulic jack, and a light disk in the thread protector 
gives further protection. 

Heads. The first “service” head, which was hur- 
riedly designed and put into production by the 
Naval Gun Factory when the high-priority service 
test was in prospect, was the 11.75-in. Rocket Head 
Mk 1 . It was admittedly a stopgap and was made 
by welding a heavy adapter to the rear of a standard 
Navy 500-lb SAP bomb and machining the buttress 
threads on the adapter. It allowed only a single 
base fuze and was not properly sealed against the 
motor pressure. Later the Mk 2 was designed hav- 
ing essentially the same shape as the Mk 1 but a 
solid nose (the Mk 1 had a small nose fuze hole) and 


tests were made using practice heads. They con- 
sisted of a piece of tubing closed at the front with a 
standard dome-shaped welding head and are shown 
in several of the photographs. 

Fuzes. Tim started out with the Mk 157 base 
fuze (Mods 1 and 2) because it was available and 
later used the improved Mk 163. The DDR fuze 
for Tim is designated Mk 162. 

Types and Designations 

The original motor (tube length 82.0 in.) was 
designated Mk 1 Mod 0 in CIT production and Mk 
1 Mod 1 in BuOrd production. Mk 1 Mods 2 and 3 
were assigned to the slightly shortened version 


11.75-IN. AIRCRAFT ROCKETS 


195 


(motor tube 75.75 in. long), and Navy production 
was changed to this design. Both rounds are des- 
ignated Mk 3 with either Mk 1 or Mk 2 heads, but 
CIT nomenclature distinguishes between the long 
Model 3 and the “medium-short” Model 4. For 
the so-called “ultra-short” motor, the designation 
Mk 2 was assigned to Navy production motors, but 
none were ever produced, and motor and round 
usually go by the CIT name, Model 5. 

1954 Launchers 


Aircraft launchers for Tiny Tim (see Figure 14) 
have been discussed in Chapter 17. Several ground- 



Figure 15. Experimental Type 61 “zero-length” 
ground launcher for Tiny Tim. 


firing launchers were also built for proof-firing the 
rounds and for possible use against caves. Most of 
the launchers were of the two-rail variety, the guide 
consisting of two long parallel pipes supported so 


that one fin rides between them. Launchers of this 
type discussed in Rocket Launchers for Surface Use 43 
are the Type 5^ proof-firing launcher, the Type 55 A 
launcher mounted on a two-wheel trailer, and the 
Type 59 “portable” launcher which sits on the 
ground on its own legs and can be carried by eight 
men. All these launchers are bulky (12 to 15 ft 
long) and cumbersome and, as in the case of the 
HVAR, do not apparently increase the accuracy 
over that of a much shorter launcher. A “zero- 
length” ground launcher, the Type 61 (see Figure 
15), was therefore designed in which the front end 
of the motor is supported on a rotating sector and 
both ends became free simultaneously after 10 in . 
of motion. 


1955 Reports 

A full discussion of the design and development 
problems of the rocket motor is given in reference 44 
from which much of the previous discussion is 
taken. The state of the ammunition at the time of 
the first high-priority program, with the X-type 
charge support and the Mk 1 head, is shown in a 
BuOrd pamphlet, 45 and later revisions show the 
production model. On proposed service uses, the 
only CIT reports are references 46 and 47. Later 
reports have all been put out by the Naval 
Ordnance Test Station, Inyokern. An illustrated 
article on drop launching is contained in reference 
48. Manufacturing and inspection problems are 
discussed in reference 49. 



Chapter 20 

SERVICE DESIGNS OF SPEN -STABILIZED ROCKETS 

By C. TT. Snyder 


201 3.5-IN. SPIN-STABILIZED 

ROCKETS [SSR] 

D aring the first two tears of the project, all 
CIT’s work was with fin-stabilized rockets. 
In this we were following the lead of the British, 
but it was undoubtedly a wise choice for fin-stabilized 
rockets involved fewer and generally simpler prob- 
lems than spinners and could therefore be developed 
and put into service use more quickly. German 
rockets, however, were almost all spin-stabilized, 
and their success (especially against our Flying 
Fortresses) coupled with the hope of obtaining 
greater accuracy and more compact projectiles led 
to the initiation in 1943 of intensive research on 
spinners by both major rocket groups in this country. 

At CIT a few rounds of experimental 4.5-in. 
spinning barrage rockets [BR] had been fired by the 
“Accuracy Committee” in the spring of 1943, but 
the first successful firing of a finless rocket was on 
the following October 13. This rocket, designated 
the 3R1 (i.e., 3.0-in., Rotating, Type 1), consisted 
of a standard 20-lb 3.5-in. Mk 1 head (solid steel 
antisubmarine aircraft rocket [AR] head), a 3.25-in. 
motor tube, and a nozzle plate held in place by a 
3.5-in. diameter threaded ring. The eight nozzles, 
each with a 0.250-in. throat diameter, were canted 
tangentially at a 16-degree angle to give right-hand 
spin. Overall, the round had a length of 22.5 in. 
(6.4 calibers) and a weight of 29.75 lb. The 2.5-lb 
cruciform grain, seated on a “button” grid, im- 
parted a velocity of approximately 550 fps. On the 
first test, both integral (i.e., bored out of a solid 
nozzle plate) and insert nozzles were tried, and, 
since both were satisfactory, the insert design was 
chosen. Since an explosive head was required, the 
Mk 1 head was quickly replaced by the 3.5-in. 
Head Mk 3, having the same weight but somewhat 
greater length. It was discovered that the dispersion 
could be significantly decreased by machining the 
outside of the head to a slightly smaller diameter 
(2.45 in.) except for about an inch at the rear, so 
that the launcher contacted the rocket only at the 
two “bourrelets,” one formed by the rear portion 


of the head and the other by the nozzle ring. This 
rocket, fired seven weeks after the first one, appears 
at a casual glance almost identical with the one 
finally designed and standardized, but actually the 
development was only beginning. 

In the ensuing months, such problems as the fol- 
lowing had to be investigated. What are the op- 
timum nozzle cant angle and the maximum quad- 
rant elevation for stable flight, and how do these 
affect one another? Where should be the center of 
mass and what should be the shape of the nose to 
give minimum dispersion or maxim um quadrant 
angle? How long can the rocket be and how fast 
can it go and still remain stable? How does dis- 
persion vary with launcher length, and what is the 
effect of malalignment, of dynamic imbalance, of 
tip-off, and of wind? To discover the answers to 
many of these questions took more than a year 
and a very considerable number of rounds. 

The original exploratory' work on spinners took 
more definite form as the result of a request by the 
Marine Corps for a spinner which might be sub- 
stituted for the 75-mm pack howitzer. For this 
application, a tubular launcher mounted on a .30- 
caliber machine gun tripod was developed, the final 
model being the CIT Type 42B or Mk 40 Mod 0. 
In comparison with the pack howitzer, the rocket 
and this launcher had a considerable advantage in 
fighter weight and consequently greater mobility, 
but, because of its higher dispersion, the rocket was 
not adopted for service use. Various other possible 
uses of the rocket were suggested at different times 
and launchers for them were tested, but by the end 
of World War II no 3.5-in. spinners had been sent 
abroad. 


Design Features 

Grain. The 2.74-in. cruciform shape was chosen 
for the initial tests because of its ready availability' 
and because it was felt that , since its inhibitor strips 
would remain in contact with the motor walls 
throughout burning, it would be less subject to 


196 


3.5-IN. SPIN-STABILIZED ROCKETS [SSR] 


197 


being fractured or thrown off center by the spin 
forces. Inhibitors were cemented to both ends of 
the grain and to the rear half of the outer ends of all 
four arms. This 50 per cent inhibiting would nor- 
mally give a progressive burning curve, but with the 
very small nozzles erosion is so severe that the 
actual burning curves are quite regressive at high 
temperatures and slightly so at medium tempera- 
tures. Although the cruciform shape, having been 



Figure 1. 3.5-in. spinner components. 

originally designed for much longer grains (Aik 13) 
gave a very low loading density (internal K only 35) , 
it fulfilled the requirements and was never changed. 
It was designated Aik 23. A few early tests were 
made with 5.0-lb cruciform grains, but, because of 
the greater length and higher velocity , these rockets 
were not stable. 

During the winter of 1914 to 1945 two other grain 
shapes were tested in standard motors. One was a 
2.5-lb ‘‘hexaform* 7 (six-legged) grain, which gave 
performance almost identical with that of the 
standard except at low temperatures where it was 
superior to the standard. The other was a 3.09-lb 
tubular grain, which also performed satisfactorily, 
giving velocities above 950 fps as compared to the 
standard 750 fps. 

Igniter. Brass can igniters containing 20 g of 
powder were used originally, but they ignited the 


grain rather slowly so that the low-temperature 
performance was not good. Thirty-five-gram plastic 
case igniters were then tried, but after a thorough 
test it was found that plastic case igniters gave 50 
per cent greater dispersion than brass can igniters, 
the explanation presumably being that pieces of 
cellulose acetate were plugging some of the nozzles, 
at least temporarily. The final solution to the 
igniter problem, in this as in most other cases, was 
a metal case igniter, the Aik 18 Alod 0 containing 
30 g of powder. In this as in all spinner igniters, the 
“false crimp 77 (see Figure 14 of Chapter 22) was 
used to reduce the impact on the propellant grain 
caused by the bursting of the closely confined igniter 
case. 

Grid. Strictly speaking, the 3.5-in. SSR does 
not have a grid, but the term has always been used 
to denote the little button on which the grain sits. 
As previously mentioned, button grids came very 
nearly being satisfactory even for the 8.5-lb Aik 13 
grain, so they were the obvious choice for the much 
shorter spinners. The original grids were 2 in. high, 
but it was quickly shown that a reduction in height 
even to in. gave no significant change in per- 
formance, and this dimension became standard. 

Xozzle Plate and Ring. Eight nozzles were orig- 
inally chosen for symmetry with respect to the 
cruciform grain. Six nozzles were found to give no 
increase in dispersion and were preferable from the 
production standpoint as well as giving less erosion 
because of their larger size. As previously men- 
tioned, insert nozzles were chosen both for lightness 
and for cheapness, and furnace brazing was found 
to be the most satisfactory method of holding them 
in the nozzle plate. Any method which holds them 
securely is apparently equally good. To assure 
accurate alignment and secure fastening, very close 
tolerances were found to be required on the nozzles 
and on the holes in the nozzle plate; the former 
were centerless-ground to an outside diameter of 
0.812 + 0.001 — 0.000 in. and the holes were 
reamed to 0.812 + 0.000 — 0.002 in. 

The first guess on nozzle cant angle was 16 de- 
grees, and, although this rather large angle probably 
improved the performance of the early rather long 
spinners which were fired at low angles, a slower 
spin was required for good high-angle flight. A cant 
angle of 12 degrees proved to be the optimum not 
only for the 3.5-in. spinner but for all the 5.0-in. 
barrage spinners as well. 

Since the rockets were to be used from automatic 





198 


SERVICE DESIGNS OF SPIN-STABILIZED ROCKETS 


launchers, contact rings were required, and the 
design finally evolved was the same as that for the 
5.0-in. spinners (shown in Figure 4). The “hot” 
contact ring was molded into a bakelite insulator 
which slipped over the rear skirt on the nozzle ring, 
which itself formed the ground terminal. Rivets 
through insulated grommets held the contact ring 
in place and made electrical contact to the igniter 
lead inside the nozzle ring. To prevent ignition 
failures it was found desirable to solder the current- 
carrying rivet to the outer contact ring. 

Since the motor tube had external threads, the 
nozzle plate seated on the end of it, and proper 
nozzle alignment could be obtained by checking the 
alignment with respect to the front surface of the 
nozzle plate and checking for squareness of the end 
of the motor tube. 

Heads and Motor Tubes. The first spinner to be 
fired, using the button grid 2 in. high, had a motor 
tube 13^8 in. long and a 3.5-in. Mk 1 head, making 
the overall length approximately 25 in. Substitution 
of the Mk 3 HE head increased the overall length 
by almost 5 in., and, although this rocket had suffi- 
cient spin to be stable in spite of its length (the cant 
angle was still 16 degrees), it would not follow a 
45-degree trajectory unless the conical nose was re- 
placed by an ogive of 4 calibers radius or more 
(8 calibers was usually used). The reason for the 
superiority of the long ogive was that it moved the 
center of pressure forward relative to the center of 
mass, thus increasing the overturning moment so 
that it could cause the rocket to follow the turning 
trajectory without exceeding the permissible yaw. a 
With only half as much spin (8-degree cant angle) 
the rocket performed well at both low and high 
angles with the conical nose, but with an 8-caliber 
ogive nose was unstable at all quadrant angles 
because the stability factor was too low. 

Reduction of the length of the grid button by 1 34 
in. gave a motor tube 12J4 in. long, and the length 
of the head was successively reduced so that the 
payload dropped from 20 to 1834 and then to 1434 
lb. The 8-caliber ogive continued to be popular for 
experimental rounds, but, when the question of a 
suitable nose fuze arose, it proved to be simpler to 
use the conical Mk 100 without changing its exterior 
contour. Proper igniter design made possible a 

a The dynamics of spinners and how the yaw causes it to 
keep aligned with the trajectory is explained in Sections 
21.5.1 and 25.5. 


further reduction in motor length leaving a mini- 
mum of space at the front end. The final motor 
tube had a length of 11% in. and had a light skin- 
cut machined on the exterior to reduce variations 
in wall thickness and consequent unbalance. 

Seals. Motor seals, both front and rear, were 
identical, except for size, with those for the 5.0-in. 
spinners (see Figure 3 of this chapter and Figure 
13 of Chapter 23), but the nozzle end seal was 
changed to that shown in Figure 14E of Chapter 23 
so that the extending edges of the seal would hold 
the round in place in the tubular aluminum launcher 
Type 37D, which, at the time World War II ended, 
was expected to go into service use. 

Fuzes. Various nose fuzes were used in the course 
of development of the 3.5-in. spinner, but all were 
relatively minor modifications of the Army M48 
fuze, as is explained in detail in Rocket Fuzes. 1 * 
This design of fuze was chosen because it was found 
that the feature of optional delay or superquick 
detonation was very effective with the rocket. 
Tests showed that with the fuze set super quick, 
ground craters were about 1 ft deep and 3 ft in 
diameter; with the fuze set delay, the rounds either 
ricocheted giving airbursts with a good fragment 
pattern 20 to 30 ft wide at low impact angles, or 
dug in at high impact angles making craters 3 ft 
deep and 4 ft in diameter in hard ground. It would 
also penetrate and detonate behind about 8 ft of 
sandbags, 3 ft of logs, % in. of mild steel, or more 
than 1 ft of concrete at normal incidence. 


Designation and Types 

Only one model of 3.5-in. spinner was standardized 
and recommended for service use. It was the 3.5-in. 
Rocket Mk 5 Mod 0, designated by CIT as the 3.5- 
in./4 Model 24A. The /4 GPSR means “approx- 
imately 4-thousand-yards-range General Purpose 
Spinning Rocket,” and the model number alone is a 
sufficient designation. It consists of the 3.25-in. 
Motor Mk 13 Mod 0 (CIT Model 6), the 3.5-in. 
Head Mk 13 Mod 0 (CIT Model 12), and the 
Nose Fuze Mk 100 Mod 0 with Auxiliary Detonat- 
ing Fuze Mk 44 Mod 2. For rounds with inert-filled 
heads, the practice was to use a model number 100 
greater than that of the explosive-loaded round, 
so that the standard round (inert) is designated 
Model 124. 


5.0-IN. SPIN-STABILIZED ROCKETS 


199 


Launchers 

The original service launcher designed for Marine 
use in place of the pack howitzer was the Mk 40 
Mod 0 shown in Figure 2. A very large variety of 
guide shapes was tried in an effort to get one which 
would be easy to manufacture and give minimum 



Figure 2. Two views of Mk 40 launcher for 3.5- 
in. spinner. 

dispersion. No single design appeared definitely 
superior to all others, but that of the Mk 40, 3 ft 
long and tubular with three internal guide rails, 
was as good as any and was adopted as standard for 
most spinner launchers, both 3.5-in. and 5.0-in. 

A number of other launchers for various purposes 
were tested, including two varieties of automatic 
launcher: a light, smooth bore, aluminum tube 
launcher for use in jungle warfare and sabotage, and 
a closed-breech launcher for replacing the 37-mm 
gun on the LYT-A1 armored amphibian tractor. 


20 2 5.0-IN. SPIN-STABILIZED ROCKETS 

The last CIT rockets which saw large-scale action 
in World War II were the 5.0-in. spinners. The 
initial test of such a rocket was on January 3, 1944. 
Its primary purpose was to provide the PT boats 
with a heavy, high-velocity weapon of sufficient 
accuracy for use against the armored and armed 
barges which the Japanese were using for supplying 
their island garrisons. It will be recalled that, late 
in 1943, the Commander Motor Torpedo Boat 
Squadrons had begun equipping his PT’s with 
launchers for the 4.5-in. BR and had had good suc- 
cess with them, but the rocket’s comparatively low 
velocity and large dispersion made it far from ideal. 

It was hoped that a velocity of at least 1 ,600 fps 
could be achieved with a 20-lb payload. The first 
tests were made with a 12-lb charge consisting of 
four tubular grains. Overall length of the rocket 
was 37.2 in. or approximately 7.4 calibers, slightly 
shorter relatively than the 3.5-in. spinner which was 
current at the time. Its head was the front part of a 
5.0-in. Mk 1 (5.0-in. AR head). Rounds were fired 
at spin velocities of approximately 100, 200, and 300 
rps with various nose shapes and weight distribu- 
tions, and none would fly stably to the end of 
burning. The variety of combinations tried was 
great enough to make it reasonably certain that a 
rocket of that length could not be stabilized without 
a considerable increase in spin velocity, and rounds 
with 400-rps spins burst at the end of burning 
because of the centrifugal force. One group of 
rounds fired with a reduced charge to give a velocity 
of approximately 1,000 fps instead of more than 
1,600 fps was just on the verge of instability, two 
out of three rounds flying stably. It thus appeared 
that it was the increase in overturning moment 
associated with the supersonic velocity that was 
causing the trouble. 

An attempt was then made to shorten the round 
and lighten it as much as possible so that approx- 
imately the same velocity could be obtained with a 
shorter and lighter grain. This rocket, with a length 
of 5.76 calibers and a velocity of more than 1,500 
fps flew perfectly and gave, on its initial firing, a 
dispersion at low quadrant angle of only 4 mils . At 
the same time a 4.2-in. cruciform grain was sub- 
stituted for the multiple-grain charge because of the 
poor static performance of the latter. Subsequent 
tests showed that the length could be somewhat 



200 


SERVICE DESIGNS OF SPIN-STABILIZED ROCKETS 


increased , and 6 .3 calibers was adopted . This rocket , 
with the addition of a Mk 100 fuze, a metal case 
igniter, and the necessary electrical contact system 
became the 5.0-in./10 GPSR Model 20 and ulti- 
mately the 5.0-in. Rocket Mk 7 Mod 0. 

As soon as it had been shown that 5.0-in. spin- 
stabilized rockets up to about 6^ calibers would fly 
and indeed would give considerably better disper- 
sions than were attainable with fin-stabilized rock- 
ets, applications for them multiplied rapidly. In 
particular, the Navy was interested in a rocket 
which would supplement the 4.5-in. barrage rocket 
and have a longer range, since offshore obstacles 
such as reefs sometimes kept the rocket-firing boats 
too far away from the beachhead to accomplish 
their purpose. (This was the case, for example, 
during part of the Saipan operation in June 1944.) 
A 5.0-in. spinner seemed to offer the best possibility 
for this application, since ranges even up to 10,000 
yd were easily obtained and their shape made them 
easily adaptable to automatic launching. 

During the summer of 1944, various other models 
of 5.0-in. spinners appeared, having either the full 
10.1 -lb propellant grain of the Model 20 (later des- 
ignated the Mk 21 grain) or one half as heavy (the 
Mk 22) . Then in the fall the Navy drew up plans for 
a rocket gunboat which was to utilize the full poten- 
tialities of the spinners. The Bureau of Ordnance 
was to develop a continuously reloadable launcher 
(the Mk 102) with remotely controlled adjustable 
elevation and train, and the gunboat, which was to 
use the LSM hull, was to be designed especially for 
mounting ten of these new launchers together with 
four mortars, one 5.0-in. gun, and various auto- 
matic weapons. 

Also as part of the plan, CIT began an integrated 
development program on barrage spinners which 
was to produce rockets with three different ranges — 
5,000, 2,500, and 1,250 yd — all having the same 
weight (about 50 lb) and the same length so that 
they would all fit the same launchers and could be 
handled and stored in the same manner. For each 
range, a variety of heads would be available: 

1. Common [Cn]. Semi -armor-piercing, with ex- 
plosive D loading and a base fuze. 

2. General purpose [GP]. A moderately thick- wall 
shell (about Yi in.) with TNT loading and nose fuze. 

3. High-capacity [HC]. A thin- wall shell (about 
34 in.) with maximum TNT loading and nose fuze. 

4. Smoke [Sm], A very light- wall shell (about }/$ 


in. thick) with either WP or FS filling, a nose fuze, 
and a tetryl burster. 

5. Chemical warfare [CW]. Similar to the smoke 
head but designed for filling with chemical agents of 
lower density (1 .43 or less) . 

6. Pyrotechnic [Py] . A light- wall shell with time 
fuze and separating charge to eject an illuminating 
flare and parachute combination. 

This ambitious program was far from complete 
by the end of World War II because, in contrast to 
the case for finners, where the principal considera- 
tion in fitting a motor to a head is the thread size, 
the necessity for keeping weight and length constant 
and still getting a maximum payload for each rocket 
meant that every new design was a completely new 
problem. Out of the total of eighteen possibilities, 
six were completed, and one, the 5.0-in./5 HCSR 
Model 34, was given a round Mark number (Mk 10 
Mod 0) and put into extensive service use. 

In October 1944, experiments in forward-firing 
spinners from aircraft were begun. As might have 
been expected, the very large wind forces to which a 
rocket launched in this manner is subjected before it 
reaches its maximum spin velocity made necessary 
still shorter rockets and higher spin velocities than 
had been satisfactory for ground firing. The devel- 
opment of a satisfactory forward-firing round re- 
quired a considerable amount of research, both 
experimental and theoretical and in particular in- 
volving the solar yaw camera. More of the details 
of this research are given in Firing of Rockets from 
Aircraft , 2 and in Field Testing of Rockets. 3 By the 
fall of 1945 when the problem was turned over to 
NOTS, Inyokern, the 5.0-in. /14 GASR Model 39A, 
having a 19-lb payload and a velocity of 1,330 fps, 
had been developed to the point where its accuracy 
was as good as the best fin-stabilized aircraft rocket 
and the general problems of aircraft spinner bal- 
listics were fairly well understood. 

Spinner Designations 

The number of 5.0-in. spinner combinations 
which existed, at least on paper, was more than 
thirty, and it would serve no useful purpose to list 
them all. Each combination was distinguished by a 
round model number, but to make the terminology 
more descriptive it became customary to include in 
the designation the general type of the round, using 
the abbreviations given for the six types listed in 


5.0-IN. SPIN-STABILIZED ROCKETS 


201 


the preceding section, and the approximate range 
in thousands of yards for ground firing at 45 de- 
degrees QE. Thus “5.0-in./10 GPSR Model 20” 
signifies that the round has a “general purpose” 
head and a range of approximately 10,000 yd. Actu- 
ally its range turned out to be greater than expected, 
10,880 yd, but the designation was not changed. 
The exception to the general rule is the 5.0-in./14 
GASR Model 39 ( general purpose aircraft rocket), 
where the 14 signifies approximate velocity in 
hundreds of feet per second, since the round is too 
stable to follow a 45-degree ground-fired trajectory. 
The two models of 5 .0-in . Rocket Mk 7 were formerly 
called high-velocity spin-stabilized rockets [HVSR]. 

2022 5.0-in. Rockets Mk 7 [HVSR] 

In discussing the design of the various spinners, 
it will be convenient to take first the 5.0-in. Rockets 
Mk 7 Mods 0 and 1 , the two high-velocity spinners 
for PT boats, and later to point out the changes 


5.0-in. Rocket Mk 7, Mod 1 (5.0-in./9 CnSR 
Model 32); 

Head Mk 8 Mod 0 (Model 30) with Mk 31 base 
fuze; 

Motor Mk 3 Mod 0 (Model 9). 

Grain. The Mk 21 Mod 0 grain has a length of 
16.2 in. and weighs 10.1 lb. It is inhibited with 
four 8J4-in. long inhibitor strips. These were orig- 
inally placed two at the rear on opposite arms of 
the cruciform grain and two at the front on the 
other arms. Static tests showed no change in per- 
formance when other patterns — all four strips at 
the front, center, or rear;, or right- and left-hand 
spiral patterns — were used, but high-temperature 
field firings with patterns having no strips at the 
rear (i.e. , all four strips at the center or the front) 
showed a considerably increased tendency for 
motor bursts. Since spiral patterns were no better, 
the original pattern was made standard because it 
was somewhat simpler. Both ends of the grain are 
also inhibited, of course. 



Figure 3. 5.0-in./10 GPSR Model 20. 


1. Fuze (Mk 100) 

2. Head (Mk7) 

3. Filler (TNT) 

4. Front seal 

5. Propellant (Mk21) 

6. Button grid 

7. Nozzle ring 

8. Igniter leads 

9. Nozzle 


10. Rear seal (obsolete design) 

11. Contact ring 

12. Insulating bushing 

13. Motor tube 

14. Igniter (Mk 17) 

15. Felt washer 

16. Fuze liner containing Mk 44 Mod 2 
auxiliary detonating fuze 

17. Fuze liner ring 




which were necessary to adapt the basic design to 
the other models. The two rockets have the fol- 
lowing components: 

5.0-in. Rocket Mk 7 Mod 0 (5.0-in./10 GPSR 
Model 20); 

Head Mk 7 Mod 0 (Model 8) with Mk 100 nose 
fuze; 

Motor Mk 3 Mod 0 (Model 4) . 


Igniter. As in the case of the 3.5-in. spinner, 
some difficulty was experienced with the closely 
confined igniters. The 55-g Mk 14 (HVAR) igniter, 
for example, fractured the grain badly in partial 
burning tests at — 10 F. A smaller false-crimped 
metal igniter, on the other hand, performed satis- 
factorily with a minimum of free space at the front 
end of the motor. The igniter adopted, Mk 17 



202 


SERVICE DESIGNS OF SPIN-STABILIZED ROCKETS 


Mod 0, is the same one used in the 3.5-in. spinner b 
and is held in a hole in the center of a 1-in. thick felt 
ring, the hole being eccentric so that the igniter 
leads can come out of the case into the space be- 
tween two arms of the grain. 

Motor Tube. On the basis of experience with the 
3 .5-in. spinner that best accuracy was obtained with 
two points of contact with the launcher, two bour- 
relets were machined on the motor tube. The NE 
8735 HVAR tubing was used, which ran consider- 
ably over its nominal %-i n * wall thickness. To 
lighten it, as well as to reduce variations in wall 
thickness which might introduce dynamic unbal- 
ance, the tubing was machined to 4.937 ± 0.005-in. 
outside diameter except near the ends where the 
bourrelets were left 4.970 + 0.000 — 0.010 in. 

Nozzle Plate. The nozzle end design is shown in 
Figure 4. A height of K in. for the button grid was 



Figure 4. Details of nozzle end of 5.0-in. spinner 
motors with cruciform grain. 

chosen on the basis of static tests as the shortest 
that gave no change in the pressure-time curves. 
Eight nozzles were chosen because this number gave 
a convenient size for machining and gave an expan- 
sion ratio of 4 with somewhat less length than six 
nozzles. The choice of 12-degree cant angle was 

b The Mark number is different because of the different lead 

length. 


relatively arbitrary and a somewhat larger angle 
might have been preferable for the flat trajectories 
in which the rocket is used, but the choice was 
made to give stable flight at 50 degrees QE. The 
electrical contact system is virtually identical with 
that of the 3.5-in. spinner. The V-shaped groove 
just ahead of the contact ring accommodates a 
spring latch to hold the round in place in launchers 
such as the trailer-mounted Type 44 or the Type 
49B PT-boat launcher (see Figure 8) . 

At the request of the Bureau of Ordnance, the 
skirt on the nozzle ring was for a time made con- 
siderably thicker than is shown in Figure 4 (0.273 
in. instead of 0.093 in.) because it was felt that the 
thin skirt would not stand the forces to which the 
continuously reloadable Mk 102 launcher would 
subject it. It was later found that such was not the 
case, and the thin nozzle rings again became 
standard. 

Heads. The “general purpose” head Mk 7 (see 
Figure 7) was made by cutting off the rear 9.75 in. 
of the Mk 1 head, welding in a J^-in. thick steel 
plate as the base closure, and threading. It was 
intended to weigh 20.0 lb with the Mk 149 nose 
fuze, having been originally designed for the high- 
velocity aircraft rocket [HVAR] but never used with 
it except for experimental tests. With the nose fuze 
Mk 100 Mod 0 and the auxiliary detonating fuze 
Mk 44 Mod 2, the head weight is almost 1 lb less. 
Against unarmored or lightly armored targets, this 
head works very well. For example, in impact at 
45-degree obliquity with J'g-in. STS armor, fuze set 
super quick, it tears a hole 2 ft in diameter. With the 
fuze set delay , high-order detonation after pene- 
tration of J^-in. mild steel plate was observed at 
0-degree and 30-degree obliquity. Its more rugged 
construction was the principal factor in the choice 
of the Mk 100 fuze over the T— 28, which would not 
stand impact with 3 ^ 2 -in. plate. 

The alternate Mk 8 head was designed for use 
against somewhat heavier armor. It uses a standard 
Mk 31 projectile base fuze. Having no hole in the 
nose and being made from heat-treated NE 8744 
steel, it functions properly against 1-in. STS armor 
at up to 45-degree obliquity. On heavier plate or at 
higher obliquities, the head broke up but the fuze 
functioned. The velocity of the rocket is great 
enough that it will punch out a disk from 13 ^-in. 
STS even though the head deforms badly and 
breaks. It is thus clear that a still more rugged 
head is justified and highly desirable for this rocket. 


5.0-IN. SPIN-STABILIZED ROCKETS 


203 


20.2.3 High-Capacity Spinners [HCSR] 

The “high-capacity” series was the only one of 
the six proposed series of 5.0-in. barrage spinners 
which CIT completed. Its members are 

5.0- in./5 HCSR Model 34 (5.0-in. Rocket Mk 10 

Mod 0); 

Motor Mk 4 Mod 0 or Mod 2 (Model 6); 

Head Mk 10 Mod 0 (Model 38); 

Grain Mk 22, 5.6 lb, 9.1 in. long. 

5.0- in./2 HCSR Model 51A; 

Motor Mk 5 Mod 2 (Model 51 A); 

Head Mk 12 Mod 5 (Model 51); 

Grain Mk 24, 3.88 lb, 6.3 in. long. 

5.0- in./l HCSR Model 50D; 

Motor Mk 6 Mod 2 (Model 50B); 

Head Mk 13 Mod 0 (Model 50B); 

Grain Mk 25, 3.1 lb, 5.0 in. long. 

The motors vary in length to fit the powder grain 
and accommodate as large as possible a payload, 
keeping the overall length 32.2 in. for all three, but 
otherwise their design is identical with that of the 
Mk 3 except in the following particulars. 

Grids. For the 5.0-in./5 the same J^-in. high 
button was used, but for the two shorter ones the 
internal K is so extremely small that a in. high 
button was found to work equally well. 

Nozzle Plates. A cant angle of 12 degrees gave 
optimum high -angle flight for all three models. 
Eight nozzles were used in the 5.0-in./5, but with 
the very small propellant weights of the other two, 
four nozzles were sufficient and, of course, cheaper. 
As originally designed, the 5.0-in./5 was stable up 
to 65 degrees QE, the 5.0-in./2 up to 60 degrees, 
and the 5.0-in./l only a little above 50 degrees. It 
was found that the addition of a 13 ^-lb weight to 
the nozzle plate, held in place by a longer stem on 
the grid button, increased the limit for the latter up 
to about 57 degrees. 

Igniters. All three use the 30-g false-crimp metal 
case igniters, the designations being Mk 20 or Mk 18 
according to the length of the wires. 

Heads. The three heads are identical except for 
length, being made in three parts — rear closure, 
body, and fuze adapter — and silver-soldered to- 
gether. To insure that the head does not extend 
radially beyond the bourrelets and strike the 
launcher guides, the body walls are made thicker 
than desired (4.95 in. OD) and machined to 4.89 in. 
OD after silver-soldering so that the exterior surface 
is concentric with the rear threads. The Mk 30 


Mod 3 nose fuze was chosen for the HCSR series. 

Also designed but not bested by the end of World 
War II was the 5.0-in./10 HCSR. By using a 
cylindrical grain with a higher loading density than 
the cruciform, the propellant weight could be in- 
creased to 9 .8 lb and thus give approximately 10 ,000- 
yd maximum range to a payload about two-thirds 
that of the 5,000-yd rocket. 

202 4 Smoke Spinners [SmSR] and 
Chemical Spinners [CWSR] 

Of the SmSR and CWSR series, only the 5,000-yd 
models were completed. They are 

5.0-in./5 SmSR Model 41 A; 

Motor Mk 4 Mod 0 (Model 6) ; 

Head Model 54A; 

Grain Mk 22, 5.6-lb cruciform. 

5.0-in./5 CWSR Model 61; 

Motor Model 61; 

Head Model 61; 

Grain 4.9-lb cylindrical three-ridge. 

The former has the same motor as the 5.0-in./5 
HCSR. The latter motor is designed after that of 
the 5.0-in./14 GASR Model 39, and the greater 
compactness of the tubular grain allows an increase 
in volume of the head filler by about 15 per cent 
over the former. 

Head designs are similar to that of the HCSR 
heads except for the thinner wall and the addition of 
a tetryl burster extending almost the full length of 
the head. To keep the centrifugal force from dis- 
placing the long slender burster tube, it is supported 
at the rear by a spider and at the front by the fuze 
adapter. 

20 . 2.5 Pyrotechnic Spinners [PySR] 

Three PySR’s were designed for three different 
illuminating flares, two having approximately 5,000- 
yd range and one approximately 4,000. The latter 
used the Mk 4 motor. None of them were tested 
thoroughly, but they appeared relatively satis- 
factory in preliminary trials. The 5.0-in./4 PySR 
Model 40 is described in Ballistic Data. 4 The CTSR 
time fuze was developed for them. 1 

20 2 6 Aircraft Spinners 

Experiments in forward-firing spinners from air- 
craft began in the fall of 1944 using the 5.0-in./10 


204 


SERVICE DESIGNS OF SPIN-STABILIZED ROCKETS 


GPSR, having a spin velocity of 250 rps and an 
overall length of 6.3 calibers. The results were 
highly unsatisfactory, the dispersion being very 
large because the rounds were unstable in flight . On 
impact they did not penetrate the ground, but 
flopped about, spinning rapidly, and in a few cases 
reaching a vertical position, nose down, spinning 
like a top. A record of the yaw in a plane per- 
pendicular to the sun’s rays, obtained by a solar 
yaw camera in the head of one of the rockets, is 
given in Figure 5, where it is apparent that the 
nutation amplitude built up to a very large value. 



Figure 5. Yaw of 5.0-in./10 GPSR forward-fired 
from aircraft (taken from yaw camera record). 
Because of too great length, the rocket is unstable. 

The first successful forward firing was done with a 
“hybrid” round consisting of the Mk 4 motor from 
the 5.0-in./5 HCSR and the Mk 7 head from the 
5.0-in./ 10 GPSR. The shorter motor gave a spin of 
only about 150 rps, but the reduction of the length 
to 5.4 calibers made the round so much more stable 
in spite of it that the dispersion immediately 
dropped to about 8 mils and the yaw camera records 
began to look like that in Figure 6. 

As a result of this success, a program of research 
on propellant grains was undertaken in an effort to 


increase the velocity of this round as much as 
possible. By eliminating the space both at the front 
and the rear of the grain to an absolute minimum, 
it was found possible to use a 10.1 -in. length of 
4.25 x 1.25-in. three-ridge tubular ballistite, weigh- 
ing almost 7.9 lb. With this grain and a change in 
nozzle cant angle from 12 degrees to 16 degrees, the 
“hybrid” round became the 5.0-in. / 14 GASR 
Model 39 A, shown disassembled in Figure 7. The 
only change in the motor was to remove the button 
grid and substitute in its place a ring grid, visible 
in the photograph, which seats in slots on the front 
face of the nozzle ring. 

It was found that stability and dispersion were 
considerably better at higher spins, and the Model 
39 A has a maximum spin velocity of 310 rps. The 
large centrifugal forces which such spin velocities 
generate makes the propellant problem a difficult 
one, especially at low temperatures where the 
powder becomes brittle. The Model 39 A is not 
considered safe below about 40 F. To remedy this 
difficulty, and also to increase the velocity still 
further if possible in the hope of making the GASR 
into an effective air-to-air weapon, research with 
internal-burning grains which fit snugly into the 
motor tube was begun by CIT and has been con- 
tinued by NOTS, Inyokern. 

20 2 7 Launchers and Service Use 

The most important launchers developed for the 
5.0-in. spinners, outside of BuOrd’s Mk 102, with 
its capacity of 30 rockets per minute continuously, 
are the CIT Type 49B PT-boat launcher (Mk 50) 
and the CIT Type 52 automatic launcher (Mk 51), 
shown in Figures 8 and 9. 

The Mk 50 launcher comes in two varieties: 
Mod 0 for starboard and Mod 1 for port. The units 
are mounted on the bow of the boat by means of a 
pedestal (not shown in the picture). They can be 
swung inboard for loading and outboard to allow 
the blast to clear the deck during firing. The eleva- 
tion is adjustable, but the train is determined by 
aiming the boat itself. The rounds are fired in pairs. 
Several hundred launchers were built by CIT and 
BuOrd, and it is reported that they proved to be 
effective, but no detailed reports of specific actions 
have been made available. 

The Mk 51 automatic is very similar to the Mk 7 
automatic for the 4.5-in. BR, is intended for the 


5.0-IN. SPIN-STABILIZED ROCKETS 


205 


/ 

\ 


017 



I 

o “ 1 UP YAW 

45 : „ * 

\ 

Figure 6. Actual yaw camera record of “hybrid” GASR forward-fired from aircraft. Increased stability 
resulting from short length causes large initial yaw to damp out rapidly. 



Figure 7. Components of 5.0-in./14 GASR Model 39. Grid (third from left) fits in slots in nozzle ring 
(left). 


206 


SERVICE DESIGNS OF SPIN-STABILIZED ROCKETS 



FIRING CABLE 


PAWL 


POWER-IN 


JUNCTION BOX 


WATERTIGHT TERMINAL 


CONTACT POINTS 


CONTACT WIRES 



CONTACT WIRE 


PAWL 


CONTACT POINT 


CONTACT STOP BLOCK 


CONTACT SUPPORT BLOCK 


Figure 8. Launcher Mk 50 Mod 0 for PT boats, showing 5.0-in./9 CnSR in place. 


5.0-IN. SPIN-STABILIZED ROCKETS 


207 


same purpose, and has the same universal appli- 
cability. They were used in the Pacific on four of a 
flotilla of twelve “interim LSM(R)’s,” so-called 
because they were built to fill in until the “ultimate 
LSM(R)’s” with their ten Mk 102 launchers and 
other automatic weapons could be put into action. 



Figure 9. Mk 51 automatic launcher with full 
load of 5.0-in./5 HCSR. 


Each ship carried 85 automatics, making its total 
firepower more than a thdusand rounds without re- 
loading. The other eight ships were equipped with 
rail launchers for the 5.0-in. AR because the pro- 
duction of spinners was not yet great enough to 
supply the contemplated very large-scale use of 
rockets. The flotilla went into action on March 26, 
1945, and continued in operation through June 15. 
The Okinawa operations probably represented the 
most varied and extensive use of rocket gunboats 
during World War II, and the spinners carried their 
full share of the load. In addition to bombarding 
the beaches themselves, the rockets were used to 
neutralize towns, knock out roads and railways, and 
fire away at whatever targets the aircraft observa- 
tion spotters assigned to them. The troop command 
to which the four “spinner” ships were assigned re- 
ported that the effectiveness of the rocket fire was 
excellent at all times. 

One other use of spinners deserves to be chron- 
icled, although its effect on the outcome of World 
War II is hardly measurable. One submarine com- 
mander attached the base plate of a Mk 51 launcher 
permanently to the topside of his craft and stored 
a launcher and a supply of rockets below. When 
the submarine surfaced at an appropriate distance 
from the target, the launcher was brought out, 
attached to the deck frame, and loaded — the whole 
process taking about three minutes. In three dif- 
ferent attacks, this submarine fired a total of 72 
rockets in shore bombardment of the island of 
Honshu, reporting that all the rounds fell within the 
target area. Though it may not have helped much, 
it surely was fun. 



PART V 

/ 

ROCKET ORDNANCE: 
THEORY, PRINCIPLES, AND DESIGN 

By E. B. Bradford 


T he introduction by C. W. Snyder to Part IV 
applies equally to Part V. In Part IV he has 
reviewed solid-fuel rockets, their components and 
their launchers, primarily from the point of view 
of their employment as weapons, with special em- 
phasis on practice as exemplified in the rockets 
developed during World War II at the California 
Institute of Technology. The basic principles of 
rocket propulsion and the war-end status of the 
theory and practice covering rocket propellants and 
interior ballistics have been ably reviewed by B. H. 
Sage, R. E. Gibson, and F. T. McClure in Parts II 
and III. 

In Part V Snyder explains in greater detail the 
principles underlying the design of rockets for effi- 
cient performance in flight. He reviews rocket bal- 


listics rather thoroughly, covers its application to 
the design of rocket propellant charges and rocket 
motors, and surveys the applications to fin stabiliza- 
tion and spin stabilization . These chapters provide 
physical explanations for the rocket behavior and 
limitations cited in Part IV. 

The general conclusions are, of course, all deriv- 
able from physical principles long known, but their 
applications to rockets, with precision enough to 
make possible the rapid and substantial improve- 
ments in performance, had to await the data, much 
of it obtained by especially devised instrumentation 
made available from the extensive programs of 
rocket design and testing during World War II. 

The emphasis in Part V is technical rather than 
military. 


209 





Chapter 2 1 

GENERAL THEORY OF ROCKET PERFORMANCE 

By C. W. Snyder 


21 1 THE MECHANISM OF PROPULSION 

T he thrust force which propels a rocket is the 
reaction to the high-velocity rearward flow of 
propellant gas a out through one or more nozzles. 

Table 1 . Ballistic quantities for fin-stabilized rockets. 


6 = angle between the horizontal and the tangent to the 
trajectory at any time. 

do = quadrant angle of elevation; angle of the launcher 
above the horizontal; (degrees, radians, or mils). 

<r = yaw oscillation distance; distance rocket travels while 
executing one complete oscillation cycle in its yaw; (ft). 
An = cross-sectional area of nozzle throat (sq in.). 

Cd = aerodynamic drag coefficient; see footnote i. 

Cn = nozzle coefficient; ratio of thrust to product of nozzle 
pressure and nozzle throat area; (dimensionless). 
c = deceleration coefficient; defined by equation (16); 
(ft" 1 ). 

db = burning distance; distance measured along trajectory 
through which rocket moves while burning; (ft). 

F = thrust; force exerted on the rocket by the action of the 
jet at any time; (lb). 

G = acceleration of the rocket; (in units of g). 
g ss acceleration of gravity; approximately 32.2 ft/sec 2 . 

M = projectile mass; mass of the rocket without propellant; 

(slugs); the weight W in lb is more often used. 
m = instantaneous propellant mass; mass of propellant grain 
at any time during burning; (slugs) . 
mo = initial propellant mass; mass of propellant grain before 
burning; (slugs); the weight wo in lb is more often used. 
P = pressure in motor chamber (assuming no pressure 
gradient); (psi). 

Pn = nozzle pressure; pressure in motor chamber measured 
just to the rear of the nozzle end of the grain ahead of 
the nozzle itself; (psi) . 
t = time (seconds). 

t b = burning time; see footnote <i; (seconds). 

V ss velocity of the rocket at any time; (fps) . 

Vo = corrected velocity of the rocket (sometimes called 
“initial velocity”); velocity at the end of burning 
assuming no gravity drop and no air drag; (fps). 

V b = burnt velocity; actual velocity of the rocket at the end 
of burning; (fps) . 

V g = effective gas velocity relative to the nozzle; defined by 
equation (4) or (6); (fps). 

W = projectile weight; weight of the rocket without pro- 
pellant; (lb). 

wo = propellant weight; weight of the grain before burning; 

(lb). 

X = horizontal range assuming impact point and firing point 
at the same elevation; (ft). 


The function of the propellant is to generate gas to 
maintain high pressure and rapid discharge over a 
period of time (the “burning time” — 0.3 to 3 sec- 
onds for the rockets of interest here) . 


Momentum-Impulse-Thrust 

Relations 


In accordance with Newton’s laws of motion, the 
forward momentum of the rocket increases during 
any time interval by an amount equal in magnitude 
to the backward momentum imparted to the gas 
ejected. Using the notation of Table 1, this fact is 
expressed as follows. In an infinitesimal interval dt 
powder of mass dm is burned and, as gas, flows out 
the nozzle with an average effective velocity V g 
relative to the nozzle. It is assumed that V g is 
the same for all masses of gas. Relative to the 
earth the gas has velocity V — V Q and hence 
momentum (V — V g )dm. The rocket’s momentum 
at any time is (M + m)V , and its change during 
the interval considered is 

+ m)V]dt = (M + m)dV + Vdm. (1) 
Hence we have 

{M + m)dV + Vdm = (V - V 0 )dm ; (2) 

dm _ dV , Q v 

~M + m ~ T g W 


By integrating (3) over the burning period, dur- 
ing which m changes from m Q to 0 and V from 0 to 
Vo , we have 


V o i M -f- mo . W + Wo 

T, = ln M = ln -If-' 


( 4 ) 


a In accordance with the established practice of rocketeers, 
we shall, for brevity, refer to the product of combustion of the 
propellant as “the gas,” even though it is a complex mixture of 
many different gases. 


A less accurate but more frequently used expres- 
sion is obtained by considering that the average 
mass of the rocket during burning is M + and 


211 


212 


GENERAL THEORY OF ROCKET PERFORMANCE 


setting equal the total momenta acquired by the 
rocket and the gas. Thus 

(M + im)V o = rrioVg, (5) 

or 

Vo _ m 0 _ wo / fi x 

V g M + im 0 W + iwo W 

Note that in either equation (4) or (6) only the 
ratios of masses and velocities are involved, so that 
any convenient units may be used. Equation (6) 
is the basic equation of rocket external ballistics. 
If V g is assumed to be known, it enables us to pre- 
dict the velocity which, in the absence of gravity 
and air resistance, will be imparted to a rocket of 
given weight by a given amount of propellant. 
Actually, it is a definition of the “effective gas 
velocity” V g , and is used to calculate that quantity 
from velocities of rockets measured in field firing. 
As we shall see, the value of V g depends upon the 
propellant used, the design of the rocket, and the 
initial temperature of the propellant. For ballistite 
it is never far from 7,000 fps, and more accurate 
guesses can be made from experience with similar 
rockets, so that equation (6) can be used to predict 
to within perhaps 5 per cent the velocity attainable 
with a rocket of proposed design. The actual veloc- 
ity of the rocket at the end of burning, denoted by 
Vb, will differ somewhat from V 0 because of the 
effects of air drag and gravity. 

It should be noted that equation (6) is exact only 
if the ratio of propellant weight to rocket weight is 
very small. The error is 0.6 per cent or less for all 
rockets now in serviced but it becomes increasingly 
less accurate as the relative weight of the propellant 
is increased and must be replaced by equation (4) . 
Thus, for a rocket consisting of 63 per cent propel- 
lant and 37 per cent metal parts, equation (4) 
gives Vo = Vg (so that the last bit of gas expelled 
before burning ceased emerges with zero velocity 
relative to the earth) whereas equation (6) gives 
Vo almost 8 per cent too low. 

Another method of evaluating V a is to hold the 
rocket stationary and measure the force it exerts on 
its supports. The relation is obtained from another 
of Newton’s laws which states that the force exerted 
on the gas (and hence its reaction on the rocket) is 


b Even for the 5.0-in. Rocket CIT Model 38, the “White 
Whizzer” (see Section 19.4.2), in which the propellant is a 
higher percentage of total weight than in any service rocket, 
the error is only 1.0 per cent. 


given by the rate of change of its momentum. The 
mass of gas outside the rocket is m Q — m and its 
momentum is (mo — m)V g at any instant. Hence 

F = J [(mo - m)V a ] = -V.%- (7) 

By integration over the burning time we obtain 



In “static firing” the thrust is measured as a func- 
tion of time and the “integrated thrust” or “im- 
pulse” fFdt is calculated from the record. The 
specific impulse c or impulse delivered per pound of 
propellant burned, is (1 /w 0 ) fFdt. Like V g , this is a 
measure of the efficiency of the rocket motor. It 
is obviously desirable to have both as high as 
possible . 

211,2 Burning Time d and Acceleration 

The principal differences between the external 
ballistics of rockets and of other artillery result from 
the disparity in the times of acceleration. A rifle 
shell is accelerated only while it is in the bore, a 
time of the order of 0.01 second, whereas a rocket is 
accelerated as long as the propellant burns — roughly 
1 second. As a result, the forces exerted on a shell 
during firing are roughly a hundred times greater 
than those experienced by rockets. Force being pro- 
portional to acceleration, the acceleration of a 
rocket is an important ballistic quantity. Its aver- 
age value is determined approximately from the 
time of acceleration (customarily called “burning 
time,” t b ) by the relation : e 

£ avg = (9) 

gh 


c Specific impulse is customarily given in pound-seconds per 
pound. Effective gas velocity, thought of as efficiency, has 
dimensions poundal-seconds per pound, which is equivalent to 
velocity. 

d Since the burning does not stop abruptly, it is necessary 
to adopt an arbitrary definition of the burning time in terms of 
the shape of the pressure- time curve. Various definitions have 
been used for various purposes, but we shall not be concerned 
in this book with the differences among them. 

e Actually V & rather than V o should be used in equation (9) 
(see Section 21.1.1), but the relation is useful only for order 
of magnitude anyhow. 


THE MECHANISM OF PROPULSION 


213 


the result being given in units of g , the acceleration 
of gravity. For the types of rockets designed by 
CIT, the upper limit of permissible acceleration 
(set by propellant strength) has been found to be 
roughly lOO^r. The burning time thus cannot in 
general be less than approximately 0.3 second per 
1,000 fps of velocity. 

Again assuming the acceleration to be a constant, 
we can calculate the “burning distance’ ’ to an 
accuracy sufficient for almost all purposes from the 
simple relation: 

d b = lV b t b . (10) 


21 1 3 Relation of Pressure to Thrust 

If we inquire into the origin of the thrust F in 
equation (8) , we enter the realm of interior ballistics. 
The burning of the propellant produces a large 
quantity of hot gases inside the motor chamber at 
an equilibrium pressure which will be denoted tem- 
porarily by P. Since this pressure pushes equally 
in all directions against the walls of the chamber, 
it would produce no resultant force except for the 
fact that, on the area A N of the nozzle throat, it 
finds nothing to push against to balance the force 
on an equal area at the front end of the motor. 
The resultant force on the rocket is thus given as 

F « PA n . (11) 

This formula requires correction because of two 
phenomena which were not considered in the fore- 
going simple discussion. First, because of the 
impedance to the gas flow from the front to the 
rear of the motor chamber, a pressure gradient 
exists, and a more exact analysis will show that it 
is the nozzle end pressure Pn which must be used 
in the formula. Second, there is an additional force 
on the rocket most of which comes from the for- 
ward component of the pressure of the expanding 
gases in the nozzle exit cone. A quantitative ex- 
planation of this additional force involves thermo- 
dynamical considerations and is relatively com- 
plicated. For practical purposes, its effect is taken 
care of by introducing into equation (11) a propor- 
tionality factor Cn, the “nozzle coefficient” or 
“thrust coefficient,” which is a function of the 
nozzle shape and the pressure. The value of the 
coefficient is known from the theory of supersonic 
jets and from experimental data and is plotted in 


Figure 1 as a function, of the “expansion ratio,” 
i.e., the ratio of nozzle fexit area to throat area. 
With these two corrections, equation (11) becomes 

F = CnPnAn. (12) 

We can now eliminate F between (8) and (12), 
obtaining (when V = 0) 

y a = —AnCn 

W o 



(see footnote c) which gives us still another relation 
for determining the effective gas velocity. By 
means of a “static-firing” apparatus / a record of the 



1 2 3 4 6 8 10 20 30 40 60 

NOZZLE EXIT AREA/ THROAT AREA 

Figure 1. Theoretical values of nozzle coefficient. 


nozzle end pressure as a function of time can be 
obtained. Measuring the area under the curve with 
a planimeter gives the value of the pressure-time 
integral, and the effective gas velocity can be cal- 
culated by equation (13) if the nozzle coefficient is 
assumed known. Alternatively, if one measures 
both pressure and thrust (as is usually done), an 
experimental value of the nozzle coefficient can be 
obtained by eliminating V g between (8) and (13) : 


C N = 


fFdt 
AnJ'P Ndt 


(14) 


A typical pressure-time curve with the calculations 
on it is shown in Figure 2. 

It was mentioned previously that the effective 
gas velocity is a measure of the efficiency of the 
rocket. Equation (13) shows that it is connected 
with interior ballistic constants through the factor 


f Reports and equipment reflecting the static-firing ex- 
perience of CIT are available at the Naval Ordnance Test 
Station, Inyokern, California. 


214 


GENERAL THEORY OF ROCKET PERFORMANCE 


Cn which depends primarily on the nozzle shape so 
that, for evaluating the efficiency of a rocket propel- 
lant charge, a quantity more meaningful for interior 
ballistics is the ratio of effective gas velocity to 
nozzle coefficient. 



Figure 2. Typical pressure-time curve for ASR 
motor. 


Area under curve = 18.02 sq in. (on the scale of the original record where 
the squares are 1 in.). 

Conversion factor: 1 sq in. = 500 X 0.05 = 25 lb-sec per sq in. 

J'PncU = 18.02 X 25 = 450.5 ~ 450 (since the record is not accurate to 
better than 1 per cent) . 

Nozzle dimensions: Throat diameter = 0.781 in. 

Throat area An = 0.479 sq in. 

Exit diameter = 1.75 in. 


Expansion ratio = 



5.0. 


From Figure 1 , for expansion ratio = 5.0, pressure = 1,500 psi, Cn = 1.50. 
Propellant weight wo = 1.50 lb. 

From equation (13) : Effective gas velocity V„ = ^ AnCn J* PNdt 


OO 9 

= ^ x 0.479 X 1.5 X 450 
= 6,940 fps. 

Impulse = CnAn J* PNdt = 1.5 X 0.479 X 450 = 324 lb-sec. 

Metal parts weight M = 61 lb. 

From equation (6): 

Corrected velocity Fo = — = G > 940 5 T^ = 168.5 fps. 

M + ?mo 61.75 


21.1.4 £fj ect 0 £ Propellant Temperature 

No mention of the variation of ballistic constants 
with temperature has yet been made. These varia- 
tions are discussed in detail in Chapter 22. Like 
most other chemical reactions, the rate of burning 
of the propellant is faster at higher temperatures, 
and hence the burning time is shorter and the 
equilibrium pressure higher. Because less of the 
heat energy of the propellant is required for warm- 
ing the rocket and more is available for pushing it, 


the effective gas velocity increases with increasing 
temperature over most of the temperature range. 
In some rockets, other factors enter at very high 
temperatures to reduce V Q again. In case tempera- 
ture gradients exist within the rocket, the tem- 
perature of the surface of the propellant grain 
appears to be the controlling one. 1 

242 THE RANGE OF ROCKETS 
21-2,1 Range in Vacuum 

In the absence of air resistance, the range of a 
projectile in free flight is given by the well-known 
expression: 

X = F ° 2 S f 2e ° , (15) 

where 0 O is the “quadrant elevation,” the angle of 
projection measured upward from the horizontal. 
In this simple form, the expression gives the hori- 
zontal distance between two points on the trajectory 
at the same elevation. Thus for a rocket, if we use 
Vb instead of Fo and the actual angle of the trajec- 
tory at the end of burning instead of 0 O , it gives the 
horizontal distance between the end of burning and 
the point on the downward trajectory at the same 
height. The total range is obviously greater than 
this by approximately twice the horizontal com- 
ponent of the burning distance. The correct expres- 
sion is complicated because of the effects of tip-off 
at the launcher and because of gravity drop during 
burning. (See Chapter 24.) 


21-22 Range in Air ; Effect of Drag 

For any but the very slowest rockets, the actual 
range is considerably less than the vacuum range 
because of the resistance of the air. The discrepancy 
is only about 3.5 per cent for the antisubmarine 
rocket, 8 but more than 45 per cent for the 5.0-in. 
HVAR. h The effect of air resistance is most easily 

8 We shall use the term antisubmarine rocket [ASR] for a 
group of rockets which are frequently called “Mousetrap 
ammunition.” Although differing slightly in details, all these 
rockets were designated 7.2-in. Rocket Mk 1 Mod 0 in the 
latest Navy nomenclature. They have 2.25-in. motors and 
velocities of 175 fps or less. (See Figure 1 of Chapter 18.) 

h The 5.0-in. high-velocity aircraft rocket [HVAR] , often called 
“Holy Moses,” has a velocity (in ground firing) of 1,360 fps 
and is the fastest fin-stabilized service rocket developed by 
CIT. It is shown in Figure 5 of Chapter 17. 


SPIN-STABILIZED ROCKETS [SSR] 


215 


introduced by means of the “deceleration coefficient” 
c, 1 defined by the equation: 

dV 

- 1 = cVK w 

For velocities up to approximately 800 fps, the re- 
sisting force offered by the air is very nearly propor- 
tional to the square of the rocket’s airspeed, so that 



500 700 900 1100 1300 1500 1700 

VELOCITY IN FPS 


Figure 3. Deceleration coefficient of 5.0-in. 
HVAR. 


c is a constant which can be fairly accurately esti- 
mated from theoretical considerations 2 or measured 
experimentally in a wind tunnel or water tunnel or 
by actual field firings. Its value for service rockets 
ranges between 1 X 10 -5 and 9 X 10 -5 ft -1 . A 
knowledge of the deceleration coefficient makes pos- 

* Also frequently used is the “aerodynamic drag coefficient’ , 
Cd , which is related to c by the formula Cd = 2 Wc/Ap where 
W is the weight in pounds, c is the deceleration coefficient in 
feet -1 , A is the maximum cross-sectional area in square feet, 
and p is the density of the medium (air or water) in pounds per 
cubic foot. Cd is dimensionless. 


sible fairly accurate range calculations for rockets 
of subsonic velocities by* the use of range tables for 
shells. Such calculations are discussed in Chap- 
ter 24. 

When the velocity of a projectile begins to ap- 
proach that of sound, the air drag becomes propor- 
tional to a higher power of the velocity than the 
second, so that, if we wish to continue to use equa- 
tion (16) as its definition, the drag coefficient c must 
be considered a function of velocity. The exact 
form of this variation depends upon many factors, 
including the density, length-diameter ratio, smooth- 
ness, and nose shape of the projectile, and is not the 
same for a typical rocket as for a typical shell. 
Consequently, for high-velocity rockets, the accu- 
rate calculation of trajectories is very much more 
difficult and uncertain. A typical curve of decelera- 
tion coefficient vs velocity, that for the 5.0-in. 
HVAR, is shown in Figure 3. 

For ground-fired rockets, the result of this varia- 
tion of c is that, even though its burnt velocity is 
well above sonic velocity, the rocket quickly slows 
down to approximately 1,000 fps, so that attaining 
a range greater than that corresponding to the 
vacuum range for 1,000 fps (approximately 10,000 
yd) is extremely difficult for short-burning-time 
rockets which are expected to carry a payload. This 
fact is illustrated in Figure 4 where approximate 
ranges are plotted as a function of initial velocity 
and deceleration coefficient. 1 


21 3 SPIN-STABILIZED ROCKETS [SSR] 


The foregoing discussion has been written in terms 
of fin-stabilized rockets (usually called “finners” for 
brevity)* but it is, for the most part, equally appli- 
cable to spin-stabilized rockets (“spinners”). Before 
considering the factors in which finners and spinners 
differ drastically from one another, we shall note 
the alterations which must be made in the equations 
of the preceding pages if they are to apply to spin- 
ners. Although rockets have been made to rotate 
by a variety of devices, including canting the fins, 

J ' Figure 4 is based on reference 3 and assumes that c 
varies in the same way for all projectiles according to the 
Gavre function (see Section 24.4.2). This approximation is 
fairly accurate for shells, which have little variation in the 
ratio of length to diameter and no fins or lugs to complicate 
the problem. Its accuracy for rockets can be estimated from 
the experimental points plotted. The value of deceleration 
coefficient quoted in each case is that for velocities well below 
sonic. 



216 


GENERAL THEORY OF ROCKET PERFORMANCE 



0 1000 2000 3000 


VELOCITY IN FPS 

Figure 4. Maximum range as a function of veloc- 
ity assuming Gavre resistance function. Values of 
c apply to subsonic velocities. 

Table 2. Ballistic quantities for spin-stabilized rockets. 

5 = angle of yaw; the angle between the axis of the rocket 
and the tangent to the trajectory. 
b e = equilibrium angle of yaw; yaw angle necessary if the 
spinner is to follow a smooth trajectory. 
rj ss nozzle cant angle. 

fjL = overturning moment coefficient; defined by equation 
( 22 ). 

v = “feet per turn”; distance traversed during one revo- 
lution. 

It = total polar moment of inertia; equal to (4f + m 0 )k 2 
before burning begins. 

K = transverse radius of gyration; referred to an axis per- 
pendicular to the rocket’s axis of symmetry and pass- 
ing through its center of mass; (ft). Total transverse 
moment of inertia of loaded round is ( M + m 0 )K 2 . 
k = polar radius of gyration; referred to the long axis of the 
rocket; (ft). 

R = nozzle circle radius; perpendicular distance between 
the nozzle axis and the rocket axis; (ft). 

S = stability factor; defined by equation (23); (dimension- 
less). 

s = spin velocity; (radians per second). 


ets in which the rotation is imparted by ejecting the 
propellant gas through a number of identical nozzles, 
arranged in a circle and each inclined symmetrically 
to the axis of the rocket by a given angle. This de- 
vice for imparting spin was the one most universally 
used by all the belligerents in World War II. 

The velocity relations for spin-stabilized rockets 
have been worked out in reference 4, and the nota- 
tion used in that report is summarized in Table 2. 
Remembering that V g was defined as the effective 
velocity of the gas relative to the nozzle, it can be 
seen that, when the rocket is rotating so that the 
gas is ejected at an angle 77 with the axis of the 
rocket, only the component of the gas’s momentum 
parallel to the axis is effective in pushing the rocket 
forward, so that the “effective gas velocity” is V g 
cos 7], and the rocket gets slightly less forward 
momentum than if it were not rotating. The cor- 
rections to equations (4), (6), (8), and (13) consist 
obviously in replacing V g by V a cos 77. 

A useful, but not quite accurate, expression for 
the angular velocity of the rocket can be derived 
by considering that the escaping gas exerts a thrust 
on each nozzle equal to its rate of change of mo- 
mentum V a {dm/dt), resolving this force into its two 
components, and applying Newton’s laws that force 
equals rate of change of linear momentum and 
torque equals rate of change of angular momentum . 
Then we have 

linear momentum: 

it dm , s dV 

V„ cos r]-^- = (M + m)-^\ (17) 

angular momentum: 

RV ° sin = lT if ( 18 > 

If it were possible to treat the combination of pro- 
jectile and propellant as a rigid body and neglect 
the fact that the mass and radius of gyration of the 
propellant is constantly changing, we could sub- 
stitute (M + m)k for I T and divide (18) by (17), 
obtaining 

d s 

R tan 77 = k (19) 


using a rifled or spiral launcher or a rotating launcher , which integrates immediately into 

and allowing the blast to impinge on plates set at an 

angle either in or behind the nozzle, we shall use the g _ RV tan 77 

term “spinner” to apply exclusively to finless rock- k 2 


(20) 



FIN STABILIZATION 


217 


This expression is strictly correct in the early part 
of burning provided that we use a value of the 
radius of gyration corresponding to the projectile 
plus propellant if the propellant rotates or corre- 
sponding to the projectile alone if the propellant 
does not rotate. It is not true, however, that the 
spin velocity continues to be proportional to the 
linear velocity; the spin increases more slowly than 
this, so that, later on in the burning, equation (20) 
always gives a value of s which is somewhat higher 
than the correct one. To derive a correct expression, 
one must know whether the propellant grain rotates 
at the same speed as the motor, at some slower 
speed, or not at all. Formulas applicable to these 
cases are discussed in reference 4. Experimental 
evidence is meager, but it appears that, for single- 
grain motors, the grain rotates almost as fast as the 
motor. The question is one of little practical im- 
portance, for the incorrect assumption that the 
angle of ejection of the gas is the same as the nozzle 
cant angle involves a considerably larger error. De- 
spite the approximations involved in its derivation, 
equation (20) is useful for design purposes to give 
an estimate of the cant angle. 

It is interesting to note that theory indicates the 
possibility of an equilibrium spin velocity 4,5 which 
cannot be exceeded by a rocket with a particular 
cant angle regardless of how high its forward velocity 
may become. The rocket could be made to spin so 
fast that the rotation would carry the nozzles side- 
ways fast enough to allow the gas to flow straight 
back out of the nozzles and impart no further spin 
to the rocket. The equilibrium spin could be ap- 
proached in practice only by a rocket with a very 
large nozzle-circle radius, and all rockets made to 
date fall far short of it. 

To the approximation within which equation (20) 
is correct, the distance which the rocket travels 
while rotating once is a constant characteristic of 
the rocket. This quantity, designated “feet per 
turn,” is given approximately by 

v = “feet per turn” = — (21) 

R tan rj 

if k and R are measured in feet. In practice, v is 
always smaller than one would calculate from this 
formula because nozzles are so short that the effec- 
tive nozzle cant angle (the angle which the ejected 
gas makes with the axis) is always somewhat smaller 
than rj and cannot be measured. Hence v, which is 


easily measured photographically, is taken as one 
of the fundamental ballistic constants. 

/ 

21 4 FIN STABILIZATION 

A long cylinder having its weight uniformly dis- 
tributed along its length is in stable equilibrium 
flying through the air only when it is aligned per- 
pendicular to its direction of motion, in which 
position its air drag is obviously very large. Since, 
in practically all rocket applications, we require 
that the projectile point in the direction of its 
motion so as to reduce air drag and land on its nose, 
it is necessary to stabilize it in this position. A cyl- 
inder flying through the air nose-on is in unstable 
equilibrium. If it acquires a slight yaw, that is, if 
the direction in which it is pointing and that in 
which it is moving begin to differ by a small angle, 
then the aerodynamic forces acting at each point 
of the surface cease to be uniformly distributed 
around the circumference. It is always possible, in 
such a case, to find a single force which, if applied 
at the proper point, will produce the same effect on 
the cylinder as the sum of all the complicated aero- 
dynamic forces distributed over the surface. The 
point of application of this hypothetical force is 
called the center of pressure . It always happens 
that, unless the mass of the cylinder is concentrated 
very close to the nose, the center of pressure is for- 
ward of the center of mass so that the torque pro- 
duced by the aerodynamic forces tends to increase 
the yaw (i.e . , it is an “overturning moment”) and 
cause the cylinder to tumble. To prevent this from 
occurring, two alternatives are available. Either 
we can arrange that the center of pressure be be- 
hind the center of mass so that the moment of the 
aerodynamic forces becomes a “righting moment,” 
or we can spin the projectile so that the overturning 
moment combined with the gyroscopic effect causes 
the axis of the projectile to rotate around the direc- 
tion of motion with a constant yaw instead of 
tumbling. 

On a shell, the aerodynamic forces always pro- 
duce an overturning moment. Some rockets, no- 
tably the 3.5-in. aircraft rocket with a solid steel 
head, have their centers of mass so far forward that 
the aerodynamic forces produce a righting moment 
even in the absence of fins, at least for large yaws 
after the propellant is consumed. In no case, how- 
ever, is this righting moment large enough to pro- 



218 


GENERAL THEORY OF ROCKET PERFORMANCE 


duce the requisite stability without the necessity of 
having fins at the rear end of the rocket. The 
presence of fins increases the aerodynamic forces on 
the rear relative to those on the nose, and thus 
larger fins move the center of pressure farther back 
and increase the stability. Stability can be ex- 
pressed quantitatively by the “eccentricity ,” defined 
as the ratio of the distance between the center of 
mass and the center of pressure to the length of the 
rocket, but it is more useful and customary to give 
the “yaw oscillation distance” a. As its name in- 
dicates, a is the distance the rocket travels while 
executing one complete oscillation from maximum 
yaw back to maximum yaw in a particular direction. 
It is discussed in greater detail in Chapter 24. We 
need merely note here that a small value of a char- 
acterizes a stable rocket and it is desirable to have 
finners as stable as possible. 

21,41 Dispersion of Finners 

The most exasperating thing about fin-stabilized 
rockets is the infrequency with which they go in the 
direction that they are aimed. Their inaccuracy 
arises primarily from the failure of the line of thrust 
of the jet to pass through the center of mass of the 
rocket. This causes the rocket to rotate during 
burning about an axis through the center of mass 
perpendicular to the trajectory, with the result that 
the thrust of the motor is changed from its initial 
direction as determined by the orientation of the 
launcher. The perpendicular distance between the 
center of mass and the line of thrust (usually 
measured in thousandths of an inch) is called 
“malalignment,” and a major portion of the effort 
in designing and manufacturing a finner is directed 
toward keeping it as small as possible. 

The malalignment may vary in magnitude and 
direction during burning, but for theoretical analysis 
it is usually assumed to be a constant. In this case, 
it tends continually to increase the yaw of the rocket 
in a particular direction, and, since the yaw changes 
the direction of the thrust, a deflection in that 
direction results. What the direction of this yaw is 
depends upon the orientation of the rocket on the 
launcher, so that the directions are randomly 
oriented, left and right orientations being equally 
probable. Thus malalignment does not change the 
center of impact of a large number of rounds, but 
introduces a dispersion about this center which is 


roughly proportional to the malalignment. It would 
be expected, and was early demonstrated experi- 
mentally, that, after burning, a rocket continues in 
the direction it had at the end of burning, and no 
further inaccuracy is introduced. 

The theory of dispersion is discussed at greater 
length in Chapter 24; its predictions are summarized 
as follows : k 

1. For relatively low-velocity rockets having 
short burning times (e.g., the ASR and BR), * 1 the 
burning distance is considerably less than half the 
yaw oscillation distance. The yaw caused by the 
malalignment, therefore, continues to increase all 
during the burning so that the deflection at the end 
of burning is approximately proportional to the 
burning time. Such rockets exhibit a marked de- 
crease in dispersion with increasing temperature 
because of the shorter burning time. If a very 
accurate rocket of this type is desired, its burning 
time must be made short enough so that a large 
fraction of the burning takes place on the launcher. 

2. For high-velocity rockets such as the forward- 
firing aircraft rockets, however, the fin size rather 
than the burning time is the most important factor 
in determining the dispersion. The reason for this is 
that the restoring torque due to the fins begins to 
become appreciable fairly early in burning and 
opposes the efforts of the malalignment to increase 
the yaw. The burning time is usually long enough 
so that the burning distance is somewhat longer 
than half the yaw oscillation distance, so that, be- 
fore the malalignment torque ceases, the rocket has 
had time to reach a maximum yaw, return to zero 
yaw, and begin to yaw in a direction opposite to 
that induced by the malalignment. In the case of 
extremely long burning times, several oscillations 
may take place during burning. In either case, the 
final deflection is considerably less than that which 
would correspond to the maximum yaw of the rock- 
et. Changes in burning time have only a minor 
effect on the dispersion. 

3. For cases intermediate between 1 and 2, it is 
necessary to apply the theory in more detail (see 
Chapter 24) . 

The reason that the small malalignment torque 

k Dispersion theory is treated in detail in reference 6 and is 
summarized in reference 7. 

1 4.5-in. barrage rockets [BR] of more than six types existed, 
three of which were assigned Mark numbers. Since their basic 
design was similar we shall refer to them simply as the BR in 
cases where the differences are not involved. See Figure 3 of 
Chapter 18. 


SPIN STABILIZATION 


219 


is able to rotate the rocket appreciably is, of course, 
that the stability of a firmer depends upon its hav- 
ing a velocity relative to the air so that an aero- 
dynamic torque exists tending to reduce the yaw. 
When starting from zero air velocity, the rockets 
are stabilized only by their launchers. If the 
rocket is headed into a high-velocity wind at the 
moment of firing, the fins are able to stabilize it 
from the beginning, and much lower dispersion re- 
sults. This is the situation in forward firing from 
aircraft, and accounts for the facts that the dis- 
persion of the same rocket air-fired is usually be- 
tween 0.5 and 0.1 of its value when ground-fired 
and that aircraft rockets are designed with large 
fins so that their stability is large. 

215 SPIN STABILIZATION 


done on the motion of shells was, of course, partially 
applicable to rockets, bht the addition of the jet 
force during burning and the much greater relative 
length of rockets introduced new and complex 
phenomena which had not been observed with shells. 
During the last three years, much progress toward 
understanding them has been made, but they still 
present one of the most extensive and potentially 
fruitful fields for further research in rocketry. For 
the details of the theory, the original papers should 
be consulted. These are summarized a little more 
fully in Chapter 25, but in the following para- 
graphs we shall attempt to understand qualita- 
tively the factors influencing a spinner’s motion in 
order to see what points are important in design, 
ignoring, for the most part, the manifold com- 
plications. 


That a projectile can be stabilized by rotation 
even though the center of pressure is ahead of the 
center of mass is a consequence of the bizarre be- 
havior of a gyroscope, which moves at right angles 
to the direction in which it is pushed. Stated more 
accurately, the rule is: if a gyroscope is rotating 
about a particular axis (vertical, say) and a torque 
is applied which tends to rotate it about an axis 
perpendicular to its spin axis (east), the result is a 
motion (precession) about the third mutually per- 
pendicular axis (north) . The directions of these axes 
are most conveniently remembered by imagining 
one’s self standing behind the rocket (a very unsafe 
place to be except in imagination) and looking along 
its axis. Then, if the rocket is spinning to the 
right (clockwise), as is assumed throughout this 
discussion, a force tending to move the nose up 
results in motion of the nose to the right; a force 
tending to move the nose to the right results in 
motion down, etc. Thus, if a rocket has a yaw of, 
say, 1 degree, the overturning moment combined 
with the gyroscopic effect leaves the magnitude of 
the yaw unchanged but causes its direction to rotate 
clockwise around the trajectory. 

The motion of a spinner is determined by the 
combination of the gyroscopic action with the vari- 
ous forces and torques which act upon it. Since 
there are four distinct types of forces and four of 
torques, the complexity of a spinner’s motion is so 
great that even now their action throughout the 
trajectory is very incompletely understood. The 
fairly extensive theoretical work which had been 


Stability Factor and 
Rocket Design 


If the rocket is moving through the air with a 
velocity V, it is subject to a torque (called the 
“aerodynamic overturning moment”) tending to 
make it tumble. As long as the velocity is less 
than about 800 fps, the magnitude of the torque is 
given by 

Overturning moment = juF 2 sin 5, (22) 

where n is the “overturning moment coefficient” 
and 8 is the yaw angle . Whether this torque will be 
able to cause the rocket to tumble depends on the 
magnitude of the gyroscopic forces, which we can 
increase to any desired value by increasing the spin 
and by making the rocket relatively shorter and 
fatter. To express this fact precisely, we have a 
quantity called the “stability factor” which gives 
the ratio of the gyroscopic to the aerodynamic 
forces and is defined by the expression: 


mk 4 s 2 

~ 4AVF 2 ’ 


(23) 


Evidently we would expect that the rocket would 
be stable if S > 1, and this is actually the case for 
shells. Because of their greater length, rockets 
require a somewhat larger value, probably about 1 .5. 

A study of the expression for S reveals several 
important facts about spinner design. Suppose we 
wish to design a rocket of a particular diameter. 


220 


GENERAL THEORY OF ROCKET PERFORMANCE 


Then the 1 ‘polar radius of gyration” k is fixed, since 
in practice it is almost impossible to change it 
much from a value of about 0.27 times the diameter. 
The spin velocity s is at our disposal, but a very 
definite upper limit on it is set by the centrifugal 
force which the motor tube and the propellant will 
stand. Now suppose that we have decided upon 
the length of our rocket as well as its caliber. Then 
the “transverse radius of gyration” K is also fixed, 
and we are left with the stability factor a function 
of s/V. As discussed in Section 21.3, this ratio, 
and hence the stability of the rocket, decreases 
steadily during the burning, so that, if the burnt 
velocity is too high, the rocket will become un- 
stable sometime before the end of burning and 
begin to gyrate wildly, lose velocity quickly because 
of the enormously increased drag, and come to 
earth with a completely unpredictable orientation 
far from the original line of fire. If we are already 
spinning the rocket as fast as we dare, the only 
alternatives are to accept a lower velocity or to 
shorten the round so that S is increased by the 
decrease in K. Evidently, then, the higher the 
velocity required, the shorter the round will have 
to be. If we wish to exceed the velocity of sound, 
the problems are aggravated by the fact that ju, 
like the deceleration coefficient, ceases to be a con- 
stant in this vicinity and increases rapidly, so that 
still stubbier rockets will be required to keep S 
above the critical value. Apparently spinners for 
aircraft forward firing must have especially rapid 
spin and short length. 

It would be interesting to know just what the 
maximum possible length of spinners is for various 
calibers, velocities, and shapes, but no specific in- 
vestigation of the point was made at CIT because it 
was not of sufficient practical utility at the time. 
It is known that supersonic 5.0-in. rockets approx- 
imately 7 calibers long become unstable near sonic 
speed, but it may be that the instability could be 
cured by higher spin if it were attainable without 
grain fracture caused by centrifugal force. No dif- 
ficulty was encountered in stabilizing 6-caliber 5.0- 
in. spinners. On the other hand, subsonic 3.5-in. 
spinners are adequately stable at 7 calibers length, 
although it is open to question whether they would 
be so at higher velocity. 

An upper as well as a lower limit to the permis- 
sible stability factor exists in most cases. If the 
trajectory is very short and relatively straight and 
the rocket is not required to change its orientation, 


it may be desirable to have a very high stability. 
For example, the 5.0-in./ 14 GASR Model 39 A, m 
designed for forward firing from aircraft, has a 
stability factor in ground firing of approximately 6, 
but rockets for barrage must have a much lower 
stability. If S is extremely large, a projectile fired 
at high angles will be so “stiff” that aerodynamic 
forces cannot turn it at all, and it will maintain its 
original orientation, landing base down. n At some 
lower value of S, the projectile will be able to turn 
rapidly enough to follow a relatively flat trajectory 
and land nose first; but, as the quadrant elevation 
is increased, it will have to turn more rapidly to fol- 
low the trajectory at its peak, and eventually a 
critical angle will be reached at which it is too 
“stiff” to follow over the peak, and instability re- 
sults. To understand this effect, let us examine in 
greater detail the mechanism by which a spinner 
keeps oriented along its trajectory. 

Consider a rocket which has been perfectly 
launched so that it is not wobbling or yawing. As 
soon as it leaves the launcher, gravity begins to act 
on it, causing the trajectory to become increasingly 
curved downward and effectively giving the rocket 
an up yaw. Since the aerodynamic moment is an 
overturning one, this yaw tends to lift the nose and 
it precesses to the right. With a yaw to the right, 
the precession moves the nose down, which is what 
is required to point it along the trajectory, and the 
rocket settles into a stable state in which it is 
yawed to the right at just the angle necessary to 
give sufficient overturning moment so that it pre- 
cesses fast enough to follow the trajectory. In an 
actual case, of course, the mallaunching, the mal- 
alignment, and the dynamic unbalance will produce 
an initial yaw and a wobble so that the rocket will 
not have the equilibrium yaw. It will, however, 
oscillate about this yaw, which is a stable position, 
as can easily be seen by considering the moments 
introduced if the yaw deviates from it by a small 
amount . 

If the rocket is spinning rapidly, it will be hard to 
turn, and the equilibrium yaw will have to be large 


m 5.0-in. /14 GASR Model 39A is the CIT designation for a 
round which was developed in the summer of 1945 but did not 
receive a Navy designation. The “14” denotes its approximate 
velocity (almost 1,400 fps) and the letters stand for “General- 
purpose Aircraft Spin -stabilized Rocket.” See reference 8 and 
Section 20.2.6 of the present volume for further details. 

n This phenomenon has been observed with shells but not, 
so far as is known, with rockets, probably because sufficiently 
stable rockets have not been fired at high enough angles. 


SPIN STABILIZATION 


221 


in order to obtain sufficient torque to turn it. The 
theoretical expression for the yaw, 


2 K 2 g cos 0 


k 2 sV( 1 - 
= mg cos 0 


V l - l/S) 
kh 


4 g cos 


M F 3 


K 2 S 
k 2 Fs 


(24) 


shows also that it is largest where the velocity is 
least, i.e., at the peak of the trajectory, both be- 
cause the curvature of the trajectory increases and 
because the aerodynamic forces are reduced. For 
high-angle fire, the peak yaw must be quite large. 
If the yaw exceeds a certain critical value (about 6 
degrees is typical), other factors enter which make 
the rocket unstable, and it begins to gyrate wildly 
or, as we usually say onomatopoetically, to “wow- 
wow.” The exact cause of this instability is com- 
plicated and not well understood; it is discussed 
in somewhat greater length in Chapter 25. 

That a spinner goes through life with its nose 
pointed to the right causes it to be deflected from 
its “normal” trajectory in two ways. The lift 
gives it a drift to the right, and the Magnus force 
pushes upward so that the range is increased beyond 
what would be expected from the velocity and the 
drag. (See Chapter 25 for definitions of these 
forces.) 


Spinner Trajectory during 
Burning 


The motion during burning is much more difficult 
to visualize than that after burning because of the 
addition of the jet force. Two factors — wind effect 
and mallaunching — are important during this pe- 
riod. The theory of wind effect has been calcu- 
lated, 9,10 and its complexity can be seen from the 
following summary. 


Wind 

Cross wind from left: Light 

Medium 

Strong 

Down-range wind: Light 

Medium 

Strong 


Deflection Produced 
Right and down 
Left and down 
Left and up 
Left and down 
Left and up 
Right and up 


The definition of a “strong” wind depends upon the 
round, its lower limit ranging from about 25 to 40 
fps. The wind effect is complicated principally 
because it is nonlinear, and its magnitude is usually 
in the range of one to several mils per foot per 


second so that* it is too large to be ignored. Because 
the burning time is so snort, gustiness in the wind 
can produce large differences in effect between 
rounds, thus introducing dispersion. Their sensi- 
tivity to wind is thus a severe limitation on the 
accuracy of spinners. 

The mallaunching effect is complicated for quite a 
different reason. Mallaunching is the name given to 
any angular acceleration about a transverse axis 
which the round acquires as a result of interactions 
with the launcher. Tip-off is the most frequently 
encountered example, tending to give the round an 
angular velocity around a horizontal axis. As the 
nose drops, the precession moves it to the left, and 
the resulting left yaw makes it continue precessing 
upward. Thus the nose moves in a spiral relative 
to the center of mass, and the rocket finishes burn- 
ing in a position below and to the left of its theo- 
retical position for no tip-off. Since the drift after 
burning is to the right, low-angle rounds land to the 
left of the range line, and high-angle rounds land 
to the right. 

If the launcher constrains the rocket from moving 
in any lateral direction, as is usually the case with 
spinners, the mallaunching need not be downward. 
Either malalignment or dynamic unbalance may 
give the round a transverse rotational velocity in 
some other direction. Theory indicates that, for a 
given degree of mallaunching, the resulting disper- 
sion is reduced by a longer launcher (i.e., by a faster 
spin at launching). However, with some launchers 
it occurs that the malalignment is approximately 
proportional to the spin velocity at launching be- 
cause of the larger forces involved, so that increased 
launcher length does not result in greater accuracy. 
For this reason, it is very difficult to determine 
what type of launcher is best, and many diverse 
types and lengths have been tried. Usually the 
choice has fallen on a simple, relatively short 
launcher for obvious tactical reasons because longer 
and more complicated ones have not given appre- 
ciably smaller dispersions. 

Special Purpose Spinners 

A further result of the complexity of spinner 
trajectories and the large number of factors in- 
fluencing them is that a spinner should be tailor- 
made to a particular purpose in order to function 
with best effect. A “Jack-of-all-trades” spinner 
would probably indeed be “master of none.” In the 


222 


GENERAL THEORY OF ROCKET PERFORMANCE 


early days of spinner development, much effort was 
expended in an attempt to work out a compromise 
round which could be used both for accurate fire 
with a flat trajectory and for barrage at relatively 
high angles. The rounds developed for barrage had 
a stability factor of about 2 and were able to follow 
over a trajectory as high as about 55 degrees to 60 
degrees. Although satisfactory for barrage, they 
could not be made more accurate than about 8 mils, 
principally because cross winds produced deflections 
of about 2 mils per fps, so large an effect as to make 


the fire control problem virtually insurmountable in 
cases where accurate fire is desired . Since the cross- 
wind effect is approximately inversely proportional 
to the stability factor, which, in turn, depends on 
the square of the spin velocity, it appears to be 
desirable to use fast-spin rockets with flatter trajec- 
tories for most applications requiring greater 
accuracy. 11 On the other hand, greater attention 
must be paid to mallaunching and dynamic balance 
for such fast-spin rounds, so that the increased 
accuracy does not come cheaply. 


Chapter 22 

DESIGN OF ROCKET PROPELLANT CHARGES 

By C . W. Snyder 


221 GENERAL REQUIREMENTS FOR A 
ROCKET CHARGE 

T he general problem of designing a propellant 
charge is a very large one, involving choice of 
the propellant composition to be used, a grain 
shape, the number of grains in the charge, a the 
method of extruding or otherwise forming the 
grain, etc., as well as the specific problems of fitting 
the charge to the motor contemplated. The general 
problems are beyond the scope of this part of the 
report. We shall take the viewpoint of the man who 
wishes to design a rocket with particular per- 
formance characteristics using a propellant of fixed 
composition — specifically ballistite, since this is the 
only propellant which has been used in quantity in 
this country — although the discussion will be gener- 
ally applicable to other propellants as well. We shall 
not be concerned with the general theoretical 
treatment of propellant performance, more com- 
plete treatments of which will be found in The 
Interior Ballistics of Rockets, 1 and in Rocket Funda- 
mentals? 

The charge designer is usually called upon to 
meet four specifications: caliber, impulse, tempera- 
ture range, and burning time. The caliber of the 
motor fixes the grain’s external diameter, which 
must be just slightly smaller than the tube’s internal 
diameter in the case of single-grain charges or, for 
multiple-grain charges, must be such that the 
grains nest properly in the tube. With a fast- 
burning propellant such as ballistite, single-grain 
charges are preferable, and multiple-grain charges 
are used only when either (1) it is not feasible to 

a The term “propellant charge” is applied to the powder 
which furnished the energy for propulsion in its final state 
ready for assembly into the rocket motor. The term “pro- 
pellant grain” signifies a single extruded or molded piece of 
propellant of whatever size (it may even be several inches in 
diameter and several feet long) either finished or unfinished. 
Thus a charge may consist of several grains, or of a single grain 
assembled with inhibitor strips or end washers, or simply of a 
single grain with nothing attached. In this chapter, we shall 
extend the customary meaning of the word “charge” to include 
not only the powder grain or grains, but also the igniter, grid, 
and other pieces intimately associated with the grain but not 
necessarily attached to it. 


make a grain as large as the motor, or (2) the 
shorter burning time obtainable with several thin- 
walled grains is desirable. Both these considera- 
tions entered in the cases of the Army’s 4.5-in. 
rockets using solvent-extruded powder (which can- 
not be made in web thicknesses exceeding 0.4 in.) 
and of the Tiny Tim (which was too big for the 
available propellant extrusion presses). Unless 
otherwise noted, we shall consider only single-grain 
motors. 

22 2 IMPULSE AND GAS VELOCITY 

The impulse or integrated thrust delivered by 
the motor is given by equation (8) of Chapter 21 
as the product of the mass of the grain by the 
effective gas velocity V g . The theoretically attain- 
able gas velocity depends on the propellant com- 
position, 2a and in practice its value for ballistite is 
approximately 7,000 fps, which is higher than that 
for most other propellants because of the larger heat 
of combustion of ballistite. For any particular 
motor, V 0 must ultimately be determined experi- 
mentally by measurements of that rocket’s velocity 
in the field, but in preliminary design calculations 
one may use a value determined from tests of a 
similar motor. 

Gas velocities for typical CIT rockets are shown 
in Figure 1 . It will be noted that V g increases with 
temperature over most of the temperature range, 
the reason being partly that, when the propellant 
and metal parts are cold, more of the heat energy 
liberated is consumed in heating them up and less 
is available as kinetic energy of the gas to impart 
momentum to the rocket. That the curve usually 
turns down again at high temperatures results from 
the decreased strength of the propellant, more and 
more of the powder being broken up and expelled 
without being burnt. The decline at high tem- 
peratures is absent when the grain is not subject 
to large forces (from acceleration or pressure drop) 
and when high-strength propellant is used, as can 
be seen from the curves. 



223 



EFFECTIVE GAS VELOCITY IN FPS 


224 


DESIGN OF ROCKET PROPELLANT CHARGES 


7,400 


7,000 

6,600 

6,200 

5.800 

7.800 

7.400 
7P00 

6,600 

6,200 

5.800 

5.400 

7.800 

7.400 

7.000 

6,600 

6,200 

7.200 

6.800 

6.400 

6.000 

5.600 

7.400 

7,000 

6.600 

6.200 

-40 -20 0 2C 40 60 80 100 120 140 160 

PROPELLANT TEMPERATURE IN DEGREES F 

Figure 1. Gas velocities for various CIT rockets. 

Effect of web thickness and acceleration, curves A, B, C, D: 

Constant: Motor inside diameter 2.0 in. 

Grain outside diameter 1.7 in. 3-ridge tubular 
Grain weight 1.43 or 1.44 lb 

Variable: Grain inside diameter, web thickness, and length 
Projectile weight and acceleration 

Curve A: 7.2-in. Rocket Mk 1 Mod 0 (antisubmarine rocket); 60.1 lb 

2.25-in. Motor Mk 3 Mod 1 
Mk 1 Mod 0 Grain, 1.70 X 0.59 X 11.6 in. 

Curve B: 4.5-in. Rocket Mk 1 Mod 0 (barrage rocket); 28.7 lb 

2.25-in. Motor Mk 7, 8, or 9 Mod 0 
Mk 1 Mod 0 Grain, 1.70 X 0.59 X 11.6 in. 

Curve C: 7.2-in. Rocket CIT Model 18 (demolition rocket); 61.8 lb 

2.25-in. Motor CIT Model 14 
Grain 1.67 X 0.90 X 14.9 in. 







Curve D: 4.5 BR (special with short burning time); 29.5 lb 

2.25- in. Motor CIT Model 7 
Grain 1.7 X 0.9 X 14.5 in. 

Remarks: Differences at low temperatures not understood. Note absence 
of acceleration effect at these low accelerations and marked 
drop in V g for thin- web grains at high temperature. 

Effect of length or internal K for thin-web grains, curves E, F, G: 

Constant: Rocket type 7.2-in. retro rocket 

Motor type 3.0-in. inside diameter with box grid 
Grain shape 2.5 X 1.4 in. 3-ridge tubular 
Projectile weight and acceleration (approximately) 

Variable: Grain length and weight and internal K 
Curve E: 7.2-in. Rocket Mk 7 Mod 0; 64.3 lb 

3.25- in. Motor Mk 1 Mod 0 

Mk 6 Mod 1 Grain, 8.7 in. long, 1.8 lb, Ki = 37 
Curve F: 7.2-in. Rocket Mk 10 Mod 0; 67.3 lb 

3.25-in. Motor Mk 2 Mod 0 

Mk 7 Mod 1 Grain, 13.0 in. long, 2.8 lb, Ki = 55 
Curve G: 7.2-in. Rocket Mk 12 Mod 0; 70.6 lb 

3.25- in. Motor Mk 3 Mod 0 

Mk 8 Mod 1 Grain, 19.6 in. long, 4.1 lb, Ki = 82 
Remarks: Note efficiency at high temperature — extremely high for short- 
est grain, extremely low for longest grain, the latter due to 
thin web. 

Effect of length or internal K for thick-web grains, curves H, I: 

Constant: Rocket type 2.25-in. subcaliber aircraft rocket 

Motor type 2.0-in. inside diameter with box grid 
Projectile weight approximately 12 lb 
Grain shape approximately 1.7 X 0.27 in. 3-ridge tubular 
Variable: Grain length and weight and internal K 
Curve H: 2.25-in. Rocket Mk 2 Mod 0 

2.25- in. Motor Mk 12 Mod 0, Kn = 225 

Mk 17 Mod 0 Grain, 8.5 in. long, 1.12 lb, Ki = 77 
Curve I: 2.25-in. Rocket Mk 1 Mod 0 

2.25-in. Motor Mk 10 Mod 0, Kn = 190 
Mk 16 Mod 0 Grain, 12.5 in. long, 1.75 lb, Ki = 121 
Remarks: Note better high-temperature performance shown in curve I 
than in curve G because of thicker web. Curve I represents 
the heaviest grain that has been successfully used in a 2.0-in. 
motor. 

Effect of propellant strength and nozzle coefficient, curves J , K, L: 

Constant: Rocket type 3.5-in. Rocket Mk 1 Mod 0 (AR) 

Grain Mk 13 Mod 0, 2.74-in. cruciform, Kn = 167 
Variable: Motor inside diameter at nozzle end of grain, internal K 
Expansion ratio of the nozzle 
Propellant composition and strength 
Curve J: 3.25-in. Motor Mk 7 Mod 0, Ki = 112 

Propellant JP; ultimate strength 270 psi at 140 F 
Expansion ratio 4.0; nozzle coefficient 1.47 
Curve K: 3.25-in. Motor Mk 6 Mod 0; Ki = 130 

Propellant JP; ultimate strength 270 psi at 140 F 
Expansion ratio 2.35; nozzle coefficient 1.40 
Curve L: Same as curve J except: 

Propellant RDS 1154.2; ultimate strength 700 psi at 140 F 
Remarks: Note improved high- temperature performance with high- 
strength propellant. Effect of Ki is not apparent in these 
curves but shows up in burst frequency at high temperature. 

Gas velocity of large motors, curves M and N, points 0 and P : 

Constant: Grain shape 4.2-in. cruciform 
Variable: Motor design and grain size 

Propellant composition and strength 
Curve M: 5.0-in. Rocket Mk 4 Mod 0 (HVAR) 

5.0-in. Motor Mk 1 Mod 0 
Mk 18 Mod 0 Grain, 24.0 lb, 39 in. long 
Propellant JPN; ultimate strength 230 psi at 140 F 
Curve N: 11.75-in. Rocket Mk 3 Mod 0 (Tim) 

11.75-in. Motor Mk 1 Mod 0 
4 Mk 19 Mod 0 grains; 36.3 lb, 60 in. long 
Propellant JPN; ultimate strength 230 psi at 140 F 
Point O: Same as curve M except: 

Propellant JPH; ultimate strength 550 psi at 140 F 
Point P: Same as curve N except: 

Propellant JPH; ultimate strength 550 psi at 140 F 


EQUILIBRIUM PRESSURE IN A MOTOR 


225 


From the viewpoint of external ballistics, V g is 
related to the velocity of the rocket [equation (6) 
of Chapter 21]; from the viewpoint of internal 
ballistics it is related to the pressure-time curve, 
which can be determined by static firing. Equation 
(13) of Chapter 21 shows that it is determined by 
the area under the pressure-time curve and has no 
relation to its shape. 

223 PRESSURE-TIME CURVES 

To determine the exact shape of the pressure-time 
curve required, we turn to the temperature-range 
specification, which usually reads “as large as 
possible,” although for certain applications either 
the upper or the lower temperature limit may be 
more critical. With ballistite (and all other propel- 
lants in common use), the pressure in the motor 
during burning is considerably increased as the 
initial propellant temperature is increased, yet the 
maximum pressure developed at any point in the 
burning must never exceed the safe working pres- 
sure of the motor assembly at the highest propellant 
temperature to be encountered in service. On the 
other hand, in order to maintain satisfactory burn- 
ing of the powder, the reaction pressure at the 
lowest firing temperature must not fall below a 
certain limit, normally about 300 psi. Thus it can 
be seen that the maximum temperature range 
over which the motor will function satisfactorily is 
obtained when the maximum reaction pressure at 
any given temperature is as low as possible and the 
minimum pressure at the same temperature is as 
high as possible. This condition is obtained when 
the pressure remains essentially constant through- 
out the reaction period, or, as we say, when the 
burning is “neutral.’ 

The situation is modified slightly by the fact 
that in many cases the motor walls are heated suffi- 
ciently by the propellant gas to decrease in strength 
toward the end of burning, and the maximum pres- 
sure which can be permitted is consequently lower 
at the end of burning than during the first part. But 
it has also been found that the minimum pressure 
at which satisfactory continuous burning can be 
maintained decreases as the motor walls are heated 
and is therefore somewhat lower during the latter 
part of burning than the pressure necessary to ignite 
the grain. These two considerations indicate that 
the most desirable pressure-time curve will usually 


be one which is slightly “regressive,” that is, one in 
which the pressure decreases somewhat from the 
start to the end ,of the reaction. The exact amount 
of regression to produce optimum results in a given 
application must usually be determined by experi- 
ment, but it is seldom large. 

224 EQUILIBRIUM PRESSURE IN A 
MOTOR 

In order to see how the shape of a pressure-time 
curve can be controlled, we shall examine the fac- 
tors which determine the motor pressure. The first 
of these is the burning characteristics of the powder. 
If any substance which does not require an exterior 
supply of oxygen to support its combustion is ignited 

Table 1. Internal ballistics quantities. 

j3 = proportionality constant relating motor pressure to 
burning rate; it is a function of temperature, 
p 5s propellant density. 

An = nozzle throat area. 

A s = surface area of the grain. 

B = linear burning rate of propellant. 

Cd = nozzle discharge coefficient. 

D = internal diameter of motor tube. 
d = external diameter of grain (excluding ridges or in- 
hibitors) . 

5 ss internal diameter of grain (cylindrical) . 

Ki = “internal K”; ratio of charge area to port area. 

Kn = “nozzle K ratio of charge area to nozzle throat area. 

L = grain length. X = 100 L/KiD. 

Mp = propellant grain mass. 
m = mass rate of discharge of gas through the nozzle. 
m' = mass rate of production of gas by the grain. 

P = pressure at any point in the motor. 

P N = “nozzle pressure”; measured just to the rear of the 
nozzle end of the grain. 
tb = burning time. 
v = volume of grain. 

W = web thickness of grain. 
w = web thickness, w = W/D. 


simultaneously over its whole surface (and if no 
complications arise, such as obstructions to the free 
flow of gas away from certain portions of the surface) , 
burning will proceed at the same rate at every 
point, so that each surface remains always parallel 
to its original position. The accuracy with which 
the burning takes place perpendicular to the surface 
is strikingly shown by partial burnings of grains, 
some of which are shown in Figures 9 and 11. If a 
small indentation (a serial number, for instance) is 
made in the surface, it will still appear — and be 
legible — although the surface may have receded 
in. or more from its original position. In this 


226 


DESIGN OF ROCKET PROPELLANT CHARGES 


type of burning, it is obvious that the amount of 
powder which disappears, and hence the amount of 
gas generated, will depend only on two factors: 
(1) the speed with which each surface recedes, i.e., 
the linear burning rate B, and (2) the instantaneous 
ignited surface area As of the grain. 

Now the linear burning rate B is a function, un- 
fortunately, of a considerable number of variables, 
including pressure, the initial temperature of the 
grain, the velocity of gas past its surface, and the 
radiation density in its vicinity. It is most critically 
dependent upon the pressure, and hence is fre- 
quently expressed by the approximate relation : la 

B = Kxooo) (1) 

where /3 is a constant for any particular temperature. 
For ballistite, the exponent n is approximately 0.7, 
and the variation of burning rate with pressure and 
temperature is as shown in Figure 2. 



Figure 2. Burning rate of JPN ballistite. 


As soon as gas begins to be generated by the 
burning of the charge, a pressure differential be- 
tween inside and outside the motor appears, and 
gas begins to flow out through the nozzle at a rate 
which is proportional to the pressure difference 
and to the nozzle area. That is, 

( 2 ) 


where An is the nozzle throat area, Pn is the 
excess pressure just inside the nozzle, and m is 
the mass of gas discharged per second. Except for 
very low pressures, such as are not usually used in 
rocketry, the proportionality factor Cd, called the 
“nozzle discharge coefficient,” is constant for a 
given powder composition. If A N is expressed in 
square inches, Pn in pounds per square inch, and 
m in pounds per second, its value for ballistite is b 


C D « 0.00065 


lb (mass) 
lb (force) X sec 


or sec -1 . 


We have already seen that the rate of generation of 
gas by the charge depends on the linear burning 
rate B and the charge area As, so we can write 

m' = pA s B, (3) 

putting the density of the propellant p in pounds 
per square inch. Since the burning rate increases 
with pressure, the pressure will rise rapidly when 
the motor is fired, until a value is reached such 
that the rate of gas flow through the nozzle is equal 
to the rate of gas generation, or 

m — m’ — C d AnPn = pA s B. (4) 

This relation enables us to solve for the equilibrium 
pressure.® 

Pn (equilibrium) = yr" ' (5) 

G d A jv 


Although the above treatment is only approximately 
correct, since we have neglected the pressure gradi- 
ent inside the motor which causes the gas to flow 
toward the nozzle d (the pressure difference be- 
tween the two ends may amount to several hundred 
pounds per square inch) , it does correctly show that, 
except for constants depending only on the propel- 
lant composition, the instantaneous equilibrium 
pressure is determined by the ratio of the ignited 
area of the charge to the area of the nozzle throat. 
This ratio is the most important design constant in 
the internal ballistics of rockets and is denoted by 

b Theoretical derivations of the value of Cd are given in 
The Interior Ballistics of Rockets, lb and in Rocket Fundamen- 
tals. 2a The nozzle discharge coefficient Cd must be distin- 
guished carefully from the aerodynamic drag coefficient in 
Chapter 21 which was denoted by the same symbol. Nor- 
mally both quantities will not appear in the same report. 

c The product pB, in units of lb/ft 2 -sec or slugs /ft 2 -sec, is 
called the burning rate in much of the literature of internal 
ballistics. 

d A more nearly exact relation is given in The Interior 
Ballistics of Rockets. 10 


m = CdAnP n, 


EQUILIBRIUM PRESSURE IN A MOTOR 


227 


the symbol K N and called “the nozzle K” or simply sides of the web, is omitted for grains which burn 
“the K” of the motor. only on the interior surface. 


22.4.1 Advantages of Tubular Grains 

This analysis shows that, unless the burning rate 
varies during the reaction, a grain will be neutral- 
burning if it is “geometrically neutral/’ that is, if 
its surface area remains constant during the reac- 
tion. The burning rate is not strictly constant, but 
pressure-time curves do follow area-time curves 
fairly closely as shown in Figure 3. A solid cylinder 
burns regressively, a tube ignited only on its inside 
surface is progressive-burning, and a tubular grain 
which burns inside and outside but not on the ends 
is, geometrically, neutral because the internal radius 
and area increase at the same rate as the external 
radius and area decrease. It is primarily because of 
this characteristic of neutral burning that tubular 
grains are used in most rockets. Evidently, if we 
wish to introduce a little regression into the burning 
of a tubular grain, we can allow one or both of the 
ends to burn so that the length decreases during the 
reaction. If we do not wish the ends to burn, we 
must prevent the propellant gas from touching them 
by cementing on a plastic “inhibitor” disk. 

Tubular grains are popular for two other reasons 
also. First, they are easy to extrude and require 
relatively little processing after extrusion; hence 
they are inexpensive. Second, they burn up com- 
pletely because their wall thickness is the same at 
every point. As the walls become thinner uniformly 
during burning, their thickness becomes zero every- 
where simultaneously. The smallest distance be- 
tween two adjacent burning surfaces is called the 
“web thickness,” and most shapes of grains have 
some portions that are thicker than the “web.” 
After the web is burned through , the charge area is 
so small that it cannot maintain the pressure re- 
quired for continuous burning, and the pieces re- 
maining, called “slivers,” either are ejected unburnt 
or smolder slowly and contribute nothing to the 
momentum of the rocket. Obviously the burning 
time of a grain, which was the last specification the 
grain designer had to meet, is given by 


where w is the web thickness. The factor 2, which 
appears because burning takes place from both 



Figure 3. Comparison of pressure-time curves 
(solid lines) and area-time curves (dashed lines) 
for: (A) cruciform grain with no arms inhibited, 
(B) cruciform grain with two arms inhibited full 
length, (C) cruciform grain with all four arms 
inhibited full length, (D) thick-web tubular 3- 
ridge grain with radial holes, (E) thick- web tubu- 
lar 3-ridge grain with rod stabilization and no ra- 
dial holes. All grains are inhibited on both ends. 


228 


DESIGN OF ROCKET PROPELLANT CHARGES 


22.4.2 Pressure Drop 

Returning to the question of what determines the 
pressure in a rocket motor, we can apply again the 
logic by which we found that the nozzle pressure is 
fixed by the constant Kn . In the conventional type 
of rocket, all the gas is discharged at the rear, and, 
if any particular mass of gas is to be urged toward 
the nozzle as it leaves the charge surface, a pressure 
gradient must exist along the length of the charge. 
The pressure at any point must depend on the rate 
at which gas is generated ahead of it (proportional 
to its distance from the front end of the charge) and 
the port area through which it can flow (which is 
usually the same for all points along the charge). 
Hence the difference in pressure between the front 
and rear ends of the grain depends on the total 
area of the grain and the port area between the 
grain and the motor tube. This space at the nozzle 
end of the grain acts like a secondary nozzle, and for 
it we define a “K” which is 

“Internal K” = Kr = char ^ e area . (7) 
port area 

As long as Ki is much less than Kn, it is of little 
importance, but, if it begins to be too large, then the 
pressure drop along the grain also becomes large, 
and troubles appear at high temperatures. 

22.4.3 Temperature Sensitivity 

Increasing the temperature of the propellant 
causes it to burn faster because the constant /3 in 
equation (1) is a function of temperature, increasing 
by about one-third between 0 F and 140 F. e If the 
burning rate increases, so also does the equilibrium 
pressure in the same proportion according to equa- 
tion (5). But this increased pressure causes a still 
further increase in burning rate, which in turn 
increases the pressure again. Thus, for a constant 
nozzle size, the variation in equilibrium pressure 
with temperature is relatively large, pressures at 
140 F being typically from 2.5 to 3.5 times those at 
0 F with ballistite. This large temperature sensi- 
tivity is one of the primary disadvantages of ballis- 
tite as a rocket fuel. Development of less tempera- 

c Values of 0 for 0 F, 70 F, and 140 F for various propellants 
are tabulated in The Interior Ballistics of Rockets . ld The 
change with temperature varies from 15 per cent for 218B 
composite propellant to 48 per cent for JP, the original bal- 
listite composition used in CIT rockets. 


ture-sensitive propellants will simplify the problems 
of rocket design greatly and make possible lighter 
motors and increased temperature range. 

225 CALCULATION OF 

MOTOR PERFORMANCE 

We have seen that the pressure in a motor de- 
pends in a complicated fashion on the initial tem- 
perature of the propellant (which determines the 
burning rate at a given pressure), the nozzle K 
(which sets the equilibrium pressure for a given 
burning rate in the absence of a pressure drop along 
the grain) , and the internal K (which causes a pres- 
sure drop along the grain and thus alters the nozzle 
K required). Because of the interrelation of these 
factors, a calculation of the equilibrium pressure in 
a motor can only be made by successive approxima- 
tions. The relation of the quantities to the basic 
thermodynamic properties of the propellant has 
been extensively investigated by the propellant sec- 
tion of the project, and formulas have been derived 
for calculating the pressure distribution in a motor 
by means of certain simplifying assumptions. le ’ 3 
Calculations based on them must usually be altered 
by experimental correction factors to take account 
of such complications as heat turbulence and heat 
loss to the surroundings, which are too difficult to 
treat analytically. Practical calculations of motor 
pressures, then, are made with semiempirical rela- 
tions based on the theory. For convenience, it is 
desirable to have the relations in graphical form. 
For ballistite, such graphs have been published in 
reference 7, and their use is illustrated in the fol- 
lowing sample calculation. Of necessity, they repre- 
sent rather ideal conditions, but for the calculation 
of maximum pressure they have been found to be 
fairly reliable. 4 

As an example of the calculations involved in 
designing a grain, let us suppose that we are re- 
quired to meet the following specifications: 

1. Motor tube 4.625-in. inside diameter; 

2. Velocity 1,000 fps with 100 lb of metal parts; 

3. Maximum front pressure 3,000 psi at 130 F; 

4. Burning time 0.70 second at 70 F. 

We shall for simplicity use a cylindrical grain in- 
hibited on the ends, even though, as discussed in 
Section 22.6, such a simple shape is seldom used. 
The alterations in the method of calculation for 
other grain shapes are relatively obvious. For the 


CALCULATION OF MOTOR PERFORMANCE 


229 


various ballistic constants, we shall take the values 
determined for JPN ballistite. 

Calculation of Powder Weight. Assuming a gas 
velocity of 7,000, we have from equation (6) of 
Chapter 21 


m 0 


MV o 100 X 1,000 
V g -\\\~ 7,000 - 500 


13.316. (8) 


Calculation of Burning Time. The burning time 
will depend on the pressure at which we decide to 
operate the motor and the web thickness. Since 
neither of these has been determined, we have in- 
sufficient data to find the burning time. If we look 
at the curves of pressure vs nozzle K and internal K 
in the range of K’s which we expect to be using, we 
will find that pressures at 70 F are slightly less 
than half those at 130 F. Hence, provisionally, we 
take a pressure of 1,375 psi. The burning rate is, 
then, 

B = = 0-651 X 1 ' 375 °' 7 = °- 814i P s (9) 

so that the web thickness is 

w = 2 Bt b = 2 X 0.70 X 0.814 = 1.14 in. (10) 

Calculation of Grain Shape . The mass of a cylin- 
drical grain is given by 

m 0 = ^(d 2 - & 2 )Lp, (11) 


is limited by the necessity' for leaving adequate port 
area for the gas which is generated on the outside 
surface, but thp maximum permissible diameter 
would have to be determined by experiment. The 
factors involved are plotted in Figure 4, including 
the grain length, the internal K for the motor as a 

220 

200 

1 80 

1 60 

I 40 
K 

1 20 

1 00 

80 

60 

40 

3.8 3.9 4.0 4.1 4.2 4.3 4.4 4.5 

OUTER DIAMETER IN INCHES 

Figure 4. Relation of length to internal K for 
typical tubular grain with no radial holes. 



where p is the density (for ballistite, approximately 
100 lb per cu ft or 0.058 lb per cu in.) , L is the length, 
and 8 and d are the inner and outer diameters re- 
spectively. In terms of the web thickness w, this is 

m 0 = TrpLw(d — w) (12) 

or 

L(d -w)=^ (13) 

TTpW 

IQ Q 

L ^ d - 1-14 > = 1.14.X 0.0 58 = 64-3- (14) 

Choosing either the outer diameter or the length , 
we can calculate the other from this equation. Since 
we would like to make the motor short in order to 
save weight, the diameter should be made as large 
as possible consistent with good performance. This 


whole and the internal K’s for the central perfora- 
tion and the outer annular channel separately. 

The internal K’s are calculated as follows: 

Inner channel: 

As = tt8L = grain area; 

Ap = ^5 2 = port area; 

K (inner) = — p* (15) 

o 

Outer channel: 

A s = 7r dL; 

A p = ^(D 2 - d 2 ); 


230 


DESIGN OF ROCKET PROPELLANT CHARGES 


where D is the internal diameter of the motor 
tube. 


ML 


ML 


K (outer) = ^ _ rf2 - 214 _ (F 
Motor as a whole: 

Ki = 


4 L(d -j- 5) 


21.4 - ( d 2 - S 2 ) 


(16) 


(17) 


The ballistically ideal grain might be thought to 
be the one in which the two internal K ’ s were equal; 
from the graph it would have an external diameter 
of 3.89 in. and a length of 23.3 in. A saving of 3 in. 


Ki 

140 120 100 80 60 0 



Figure 5. Graph of front pressure vs nozzle K 
for 130 F. Time = 0.02 second. 


performance of grains of a given weight with various 
lengths, diameters, and web thicknesses are sum- 
marized in The Interior Ballistics of Rockets, 11 
reference 6. 

For the purpose of the present example, let us 
assume that we have decided on a grain 19.6 in. 
long, since it is clear from the graph that this is 
close to the shortest possible grain that will work. 
This choice determines the dimensions of the grain 
to be 4.42 in. by 2.14 in. and gives internal K of 
80, which is convenient for calculation. 

It remains to calculate the nozzle throat diameter 
required to give a front pressure of 3,000 psi at 
130 F. From the graph in Figure 5, f a nozzle K of 
202 corresponds to this pressure when Ki = 80. 

Since the grain’s burning area is 

A (S =7rL(d+5) = 19.67r(4.42+2.14)=404 sq in., (18) 
the nozzle throat area is 

A n = -jT- = ™ = 2.0 sq in. (19) 



100 120 140 160 180 200 220 240 

K N 


or more in length can be made, however, if the 
resulting high value of the internal A for the annular 
channel does not cause trouble, as in fact it does 
not. 5 The calculation of K (outer) assumes that all 
the gas moves toward the rear of the motor. Actu- 
ally, of course, if the K (outer) is much higher than 
K (inner), the gas generated near the front of the 
outer surface will move forward and escape through 
the central perforation, thus effectively lowering the 
K (outer) and increasing the K (inner). Ki is, 
therefore, the important design parameter. The 


Figure 6. Graph of pressure difference between 
front and nozzle ends vs nozzle K for 130 F. 
Time = 0.02 second. 

This requires a single nozzle 1.6 in. in diameter or 
multiple nozzles of correspondingly smaller size. 

Further information about the grain’s performance 
can be gotten from the graphs of Figures 6 and 7 and 

f Figure 5 is taken from reference 7, which gives similar 
curves for front pressure, nozzle pressure, and pressure dif- 
ference as functions of Kn and of temperature. 


l L 


TUBULAR GRAINS OF FAST-BURNING PROPELLANT 


231 


also from reference 7. For example, the pressure 
drop along the grain at 130 F is 225 psi, giving a 
total force on the grain 

F = AP X A m d = AP X ^(d 2 - **) 

= 2,640 lb. (20) 

Front end pressures will be slightly higher than 
those read from Figure 7 for K N — 200 and Ki = 



TEMPERATURE IN DEGREES F 


Figure 7. Graph of front pressure vs tempera- 
ture for K n = 200. 

80; at 70 and 10 F, they are approximately 1 ,350 
and 760, respectively. The former value is close 
to that assumed in calculating the burning time at 
70 F, so that calculation was correct. 

22 6 TUBULAR GRAINS OF 

FAST-BURNING PROPELLANT 

Despite the theoretical advantage of the cylin- 
drical shape, it is preferable in practice to depart 
slightly from it in the interest of loading con- 
venience. The first rockets developed by CIT had 
cylindrical grains which were held in the center of 
the motor tube by plastic tabs cemented to the 
grain. The extra operation involved in attaching 
these tabs can be eliminated by extruding the pow- 
der with three ridges extending beyond the cylin- 
drical surface and fitting closely the inside diam- 


eter of the motor tube. Most CIT rockets use 
grains of this shape. 

If a tubular grain of ballistite is fired statically 
so that a record of its pressure-time history can be 
obtained, the curve typically looks like the solid 
line in Figure 8, with a marked pressure peak near 
the middle of burning. Numerous experiments lg ’ 8 
have demonstrated that the irregularity of the 
curves is caused by unstable burning in the central 
perforation which causes very large stresses on the 
grain, and that it can be eliminated by at least 
three devices: 

1. Drilling a sufficient number of radial holes 
through the grain joining the central perforation 
with the outside (see Figure 9); 

2. Inserting in the central perforation a loose- 
fitting rod of some material which will remain in 
place throughout the burning; 

3. Making the central perforation irregular in- 
stead of circular in cross section. 9 

Of these alternatives, only radial holes have been 
used in service rockets. Rod stabilization has been 
avoided because it introduces more complexities 
into the design and the loading than do the radial 
holes. Noncircular perforations upset the balance 
between the rates at which the internal area in- 
creases and the external area decreases, and hence 
result in less satisfactory pressure-time curves, a 
disadvantage which is not serious. 

Virtually nothing is' known of the mechanism by 
which these three devices stabilize the burning. 
Consequently, if one is designing a grain with radial 
holes, he must determine the optimum number, size, 
and spacing by trial and error. Numerous experi- 
ments at CIT have yielded a number of general 
rules which are summarized as follows: 18 

1. Two or more holes in a given plane at right 
angles to the axis of the grain have no more stabiliz- 
ing effect than a single hole at the same point. 

2. The stabilizing effect of a hole is slightly de- 
creased if the diameter of the hole is made very 
small in comparison with that of the axial perfora- 
tion. However, increasing the diameter of the 
radial hole beyond 0.4 times that of the axial per- 
foration does not add to the stabilizing effect. 

3. Critical spacing between radial holes appears 
to be nearly independent of web thickness and 
diameter of axial perforation, but is a function of 
powder composition and increases as heat of explo- 
sion and burning rate decrease. Thus for ballistite, 
the “hottest” powder in general use, the maximum 


232 


DESIGN OF ROCKET PROPELLANT CHARGES 


permissible spacing is about 1 in., while for the 
slow-burning German propellant it is at least 10 in. 
so that many grains require no radial holes at all. 

4. The exact arrangement and position of the 
holes is of no importance, except that the stabilizing 
effect extends only over the region of the radial 
holes, which must therefore be distributed along the 
whole length of the grain. 

If the web thickness is greater than the radius of 
the hole, as is always the case, the effect of a radial 


seen by consideration of an example, the 2.5-in. by 
0.4-in. grain once proposed for the 3.25-in. Alt 
motor. 


Outside diameter 
Inside diameter 
Web thickness 
Radial holes 
Length 

Area without radial holes 
Initial area with radial holes 
Final area with radial holes 
(Final area) /(initial area) 


2.5 in. 

0.4 in. 

1.05 in. 

19 holes 54-in. diameter 

20 in. 

182 sq in. 

298 sq in. 

137 sq in. 

46 per cent 



O 0.05 0.10 0.15 0.20 0.2 5 0.30 0.35 0.40 0.4 5 0.50 0.55 0.60 

TIME IN SECONDS 


Figure 8. Effect of number of radial holes upon performance of tubular 3-ridge grain (1.7 x 0.25-in., 
11.4-in. long). 


hole is to increase the surface area at the beginning 
of burning, since the area of the sides of the little 
cylinder of powder removed is greater than that of 
the ends. At the end of burning, the situation is 
reversed, and the addition of radial holes decreases 
the surface area. Thus radial holes introduce a 
regression into an otherwise neutral-burning tubular 
grain. With the web thicknesses which have 
usually been used, this effect is not objectionable 
because it is partially compensated by the tendency 
of the burning rate to increase as burning proceeds, la 
so that the resultant regression is slight. The effect 
becomes large with thick- webbed grains, as is easily 


At the end of burning, the %-im holes have in- 
creased to 1.2 in. in diameter, and nineteen such 
holes in a 20-in. grain seriously reduce its ability to 
withstand the acceleration forces. 

Tubular grains with radial holes have not, in 
fact, been used where thick web has been required. 
A typical tubular grain is the Mk 1 , which has been 
used in the 4.5-in. BR. It is a three-ridge tube 1.7 
in. by 0.6 in. in diameter and 11.5 in. long, having 
twelve 34- in. radial holes and neither end inhibited. 
Because of the ridges, radial holes, and uninhibited 
ends, the burning area decreases from an initial 
value of 98.9 sq in. to a final value of 66.4 sq in., or 


TUBULAR GRAINS OF FAST-BURNING PROPELLANT 


233 







Figure 9. Partially burned tubular grains showing stabilizing effect of radial holes. At the top is a 3-ridge 
undrilled grain in which a large fissure has opened up, causing a sudden increase in burning area and a pres- 
sure peak such as is shown in Figure 8. Below are sections of two drilled grains showing that the stabiliz- 
ing effect of the holes extends only to their immediate vicinity. 


33 per cent. Of this, 7 per cent is due to the effect as the reaction proceeds introduces a compensation 
of the ridges, 17 per cent to the radial holes, and of about 26 per cent, however, so that the pressure- 
9 per cent to the ends. The change of burning rate time curves are only very slightly regressive. 


234 


DESIGN OF ROCKET PROPELLANT CHARGES 


22.7 OTHER GRAIN SHAPES 

Various shapes other than tubular have been 
used for propellant grains. 8 These fall generally 
into four categories: internal-burning, end-burning, 
external-burning, and multi web grains. Among the 
external-burning grains (which, to date, have been 
the most important) are included slab-shaped, cru- 
ciform, and three-, six-, and eight-armed grains. 
The cruciform is the only one of these now in general 
use, and, since it illustrates the problems encountered 
with any exterior-burning grain, we shall confine 
our discussion to it. CIT experience with the other 
shapes h is summarized in The Interior Ballistics of 
Rockets . lf 

22 7 1 Cruciform Grains 

In any external-burning grain, the area decreases 
steadily throughout burning, and it is necessary to 
inhibit the burning of certain portions of the grain 
if an even approximately neutral burning is to be 
achieved. The inhibiting process, although costly, 
time-consuming, and generally a nuisance, has the 
advantage of providing a large measure of control 
over the shape of the burning curve. Thus on the 
cruciform grain, naturally regressive when unin- 
hibited, a very progressive burning curve can be 
achieved (see Figure 3) by inhibiting the outer cylin- 
drical surface at the ends of the arms along their 
full length. A neutral burning curve is obtained 
when approximately 45 per cent of the curved 
surface of the arms is inhibited in the proper way. 1 

Cruciform grains have been used in preference 
to tubular (1) when large powder weights have 
been required as in the aircraft rockets, (2) in spin- 
ners where the fact that the inhibited portion re- 
mains in contact with the motor wall throughout 
burning makes them better able to withstand the 
centrifugal force, and (3) in cases where longer 
burning times were desired than available cylindrical 
shapes would provide. 

g Dimensions of most of the shapes extruded by CIT may 
be found in reference 10. Complete information on all grains 
recommended for service use is given in reference 11. 

h See also the following reports: on 2.74-in. cruciform: ref- 
erences 12 and 13; on 4.2-in. cruciform: reference 14; on 
hexaform: reference 15; on triform: reference 16. 

* For reasons which are not understood, the pattern of the 
inhibiting strips is critical. The effect is probably akin to that 
which causes unstable burning in tubular grains without 
radial holes. Experiments with various inhibiting patterns 
are summarized in The Interior Ballistics of Rockets. 11 


22.7.2 Internal-Burning Grains 

A tubular grain which burns only in the central 
perforation is normally very progressive but can be 
rendered neutral by putting longitudinal grooves in 
the central perforation so that it is roughly gear- 
shaped in cross section. Research on internal- 
burning grains has been very limited until the last 
few months, but they hold considerable promise 
because of the high loading density which they 
provide and because of the elimination of problems 
associated with heating of the motor tubes. The 
inhibiting problem is considerably more severe here 
than with external-burning grains because of the 
large surface area which must be inhibited. If the 
propellant can be molded instead of extruded, a 
good way to make a charge is to mold it in the motor 
tube in direct contact with the tube walls. 


22-7,3 End-Burning Grains 

End-burning grains are similar to internal-burning 
grains in their design problems. They can be used 
when extremely long burning times and relatively 
small thrusts are required. CIT experience with 
them is summarized in The Interior Ballistics of 
Rockets . lh 

2274 Multiweb 

Two types of multi web charges were investigated 
in the early days: two concentric tubular grains and 
the “4-spoke” or “okra” grains. They were designed 
primarily to get shorter burning times without sac- 
rificing loading density. No service requirement 
materialized for such charges and no summary of 
CIT experience with them exists. They are dis- 
cussed in a number of reports, however. 17,18 


22 8 LOW-TEMPERATURE PERFORMANCE 

As the temperature is decreased, the effective 
gas velocity of a rocket motor decreases, the reaction 
pressure decreases, and the burning time becomes 
correspondingly longer. In some applications these 
factors may so decrease the accuracy or range that 


IL 


MOTOR FAILURES AT HIGH TEMPERATURE 


235 


the rocket ceases to be tactically useful, but ordi- 
narily the specified lower temperature limit is deter- 
mined by the temperature at which the motor 
ceases to burn continuously. 

In the open air, a stick of ballistite will, of course, 
burn at atmospheric pressure. Inside a motor, 
however, where no oxygen is present after the pro- 
pellant gas has swept the air out of the chamber, 
the chemical reaction involved in combustion is 
somewhat different and will not proceed if the pres- 
sure drops below a minimum which depends on the 
propellant composition but is usually several hun- 
dred pounds per square inch. At lower pressures, 
the rate of the reaction is so low and the transfer of 
heat from the gas to the solid grain is so poor that 
the grain surface is not kept warm enough and 
burning may stop. When this happens, the par- 
tially burned grain can sometimes be recovered, 
but more often the motor walls, grid, and other 
metal parts in contact with the grain have been 
heated enough that they reignite the grain, which 
burns more or less normally for another period and 
perhaps again goes out. This process is usually 
called “chuffing” and motors have been known to 
“chuff” as many as fifteen times. The time between 
successive chuffs may vary between half a second 
and several seconds, and in rare instances periods of 
more than a minute have elapsed between the first 
burning period and the first chuff. Chuffing is a 
serious matter not only because it will cause the 
rocket to miss its target completely, but especially 
because the first period of thrust may be just suffi- 
cient to free the rocket from the launcher and subse- 
quent chuffs may send it in an unpredictable direc- 
tion with nearly its normal velocity so that the 
fuzes may be armed. 

Intermittent burning of this type is caused pri- 
marily by too low a pressure. Chuffing is associated 
with low-temperature firing only because sufficiently 
low pressures are not otherwise obtained normally. 
Hence the lower temperature limit can be made as 
low as desired by operating the motor at a high 
nozzle K so that the pressure is kept up. This can 
usually be done only at the expense of high-tem- 
perature performance, so that the choice of nozzle K 
depends partly on the relative importance of the 
two ends of the temperature scale. By the use of a 
blowout disk (see Chapter 23), it is possible to 
operate at high K for low temperatures and low K 
for high temperatures and thus extend the working- 
temperature range at both ends. 


22 9 MOTOR FAILURES AT 

HIGH TEMPERATURE 

22,9,1 Types of Failures 

Failures of nonrotating motors at high tem- 
peratures are of three principal types. 

1. On low-performance motors where the grain 
is subjected only to small forces, failure may occur 
because the normal operating pressure of the motor 
at that temperature is too high for the strength of 
the metal parts. Since motors are usually designed 
with a safety factor of 1.5 or 2 at 120 F to 130 F, 
the temperature required for this type of failure is 
high — perhaps 160 F to 170 F for ballistite. Failure 
occurs either at the weakest part of the motor (for 
example, the nozzle is frequently ejected) or at the 
front end of the tube where the pressure is highest. 
It takes place very early in burning before the 
rocket leaves the launcher. 

2. On long-burning-time high-performance rock- 
ets, the motor tube may be so weakened by heat 
that it will burst at relatively low pressure, opening 
up just ahead of the nozzle, where the heating 
effect is greatest. Such bursts are not very violent 
and can only occur at almost the end of burning. 

3. Most frequently, motor failures result from 
collapse of the grain, the sudden increase in burning 
area sending the pressure skyrocketing. Such bursts 
are usually extremely violent, at times amounting 
almost to a detonation, and occur sometimes imme- 
diately upon ignition, imparting almost no velocity 
to the head, sometimes a few feet off the launcher, 
and sometimes near the end of burning when the 
web has become thin. Whether the front end or 
the nozzle end of the tube fails seems to depend on 
the motor. In many cases the tube opens up along 
its whole length . Grain failure is the cause of almost 
all bursts of high-performance motors and a con- 
siderable proportion of those of low-performance 
motors. With some powders which are exceptionally 
brittle when cold, grain failures may occur at low 
temperatures. In the category of grain failures are 
included also the bursts of spinners near the end of 
burning because of fracture of the propellant by the 
centrifugal force. 

In addition, motor bursts may result from faulty 
design, such as insufficient radial holes, incorrect in- 
hibiting pattern, or insufficient space at the front 
end of the motor so that the igniter fractures the 
grain. We shall consider only the failures of rea- 
sonably well-designed motors. 


236 


DESIGN OF ROCKET PROPELLANT CHARGES 


22 9 2 Stresses on Grains 

During burning, the propellant grain is subjected 
to longitudinal compressive stresses from the fol- 
lowing sources: (1) difference in pressure between 



Figure 10. Compressive stress in Mk 13 cruci- 
form grain, 50 per cent burnt. 

front and nozzle ends of the grain, (2) skin friction 
between the flowing gas and the grain, (3) impact 
of the flowing gas on projecting portions of the 
grain, and (4) acceleration of the rocket. All these 
types of stresses increase with increasing tem- 
perature, types (1) and (4) very markedly. In 


Table 2. Total compressive stress acting on Mk 13 
propellant grain in firing at 140 F. 


Per cent 

Total stress (psi) 

of grain 



burned 

Static 

Flight 

0 

323 

441 

10 

246 

364 

29 

164 

282 

48 

119 

237 

67 

96 

214 

86 

90 

208 

91 

101 

219 

95 

127 

245 


motors of orthodox design, having the grain sup- 
ported at the rear end and all the gas flowing toward 
the rear, all the forces act in the same direction, 
and the maximum stress in the grain occurs at the 
nozzle end. The forces are discussed in greater 
detail in The Interior Ballistics of Rockets , 1 from 
which Table 2 and Figure 10 are taken, and in 


reference 19. These figures represent conditions 
for the Mk 13 grain used in the 3.25-in. AR motor. 
The discontinuities in the curve of Figure 10 are 
due to the localized impact forces on projecting 
portions of the grain at the front end of the in- 
hibitor strips (see Figure 11). It is seen that the 
most important forces are the pressure differential 
and the acceleration. To reduce the former, it is 
necessary to reduce K I} but, for a given propellant 
shape, Kj is proportional to the grain weight and in 
high-performance motors is necessarily large. Un- 
fortunately, Ki increases slightly with increasing 
temperature because the powder has a larger coeffi- 
cient of expansion than the motor. High accelera- 
tion also is usually associated with high-performance 
motors, so that obtaining good performance at high 
temperatures is the most difficult problem in design- 
ing such a motor. 

22.9.3 Mechanism of Failure 

The mechanism by which grain failures occur is 
discussed in The Interior Ballistics of Rockets . H 
When burning starts, the compressive stresses on 
the grain cause it to become shorter and fatter, 
bulging particularly at the nozzle end where the 
stresses are greatest. The bulging decreases the 
port area around the grain and hence causes the 
pressure drop to become still larger. The amount 
of bulging is determined by the elastic modulus and 
Poisson’s ratio for the propellant. A “strong” grain 
will bulge only slightly and equilibrium will be 
reached at a higher pressure than normal because of 
the higher iv/; j but a “weak” propellant will bulge 
so much that an unstable condition results, higher 
front end pressure causing more bulging which in 
turn causes still higher pressure. Thus the effect is 
virtually as if the nozzle were closed, and the pres- 
sure quickly builds up to a value which will fracture 
the grain and burst the motor. It can be seen from 
this analysis that the ultimate compressive strength 
of the propellant is of secondary importance, but 
the elastic modulus and Poisson’s ratio primarily 
determine the minimum value of the pressure drop 
at which instability begins. k 

At the end of burning, a grain fails because it 
becomes too slender relative to its length to with- 

3 There is some evidence that oscillations of the grain about 
its new equilibrium shape can occur. 20 

k Tests of the strength of various propellants are summarized 
in references 21-24. 


DESIGN OF MAXIMUM WEIGHT GRAINS 


237 


stand the forces — principally the acceleration force 
— and buckling occurs. Usually the collapse does 
not burst the motor but merely puts a sharp peak 
at the end of the pressure-time curve. Increasing 
the acceleration moves the peak closer to the begin- 
ning of burning so that more powder is ejected un- 
burnt and the effective gas velocity decreases. One 
important reason for the better performance of 
cruciform than tubular grains in high-performance 
motors is the fact that the inhibited portions of the 


of the burning surface, the burning rate in these 
regions is considerably accelerated, probably be- 
cause such conditions are conducive to more rapid 
transfer of heat from the gas to the grain. This 
effect is termed “erosive burning. ,, Since the gas 
velocity is always higher at the nozzle end, the 
burning tends to be faster at this end, but the effect 
is compensated by the pressure dependence of the 
burning rate which tends to make burning faster at 
the front. If the ratio Ki/Kn is too high, however, 



Figure 11. Mk 13 cruciform grain, inhibited. 


former continue to be supported by the motor walls 
throughout burning. 

In spinner motors, the centrifugal forces can also 
induce grain fracture and burst the tube, and it is 
frequently this consideration which limits the spin 
velocity. Two types of failures have been observed. 
At high temperatures the spin is increased because 
of the increase in effective gas velocity, and bursts 
occur near the end of burning at a temperature 
which depends principally on the tensile strength of 
the propellant. Bursts may also occur at very low 
temperatures, and here they seem to be associated 
not with low tensile strength but with brittleness of 
the propellant. 

22 10 DESIGN OF MAXIMUM WEIGHT 
GRAINS 

It is frequently required to design a motor which 
contains the maximum possible amount of propel- 
lant for the given caliber either with a specified web 
thickness or regardless of web thickness. The most 
important factor limiting the powder weight is the 
internal K. We have already seen that large K T 
means large forces on the grain, but another effect 
is involved which is perhaps equally important. 
Whenever the gas flows at high velocity over part 


the erosion becomes very severe so that the de- 
creased web thickness at the nozzle end contributes 
to the grain collapse. In practice it has been found 
that values of Ki/Kn greater than about 0.75 
cannot be used, and for service rockets it has not 
been felt desirable to exceed 0.60. For CIT rockets, 
which have Kn usually in the neighborhood of 200, 
this sets the upper limit of K r at 120. Only on the 
3.25-in. Motor Mk 6 was this value exceeded, and 
this was one of the reasons for abandoning it in 
favor of the Mk 7. 

With the stipulation that the internal K cannot 
exceed a certain value, it is simple to plot curves of 
the relations between external diameter, web thick- 
ness, length, and Ki from which the maximum 
possible tubular grain for a given motor caliber can 
be obtained. The radial holes and supporting ridges 
complicate the calculations considerably and hence 
are usually omitted from the calculations, so that 
the curves can be used to show qualitative relations 
only. A series of such curves showing the effect on 
the internal K of a constant weight grain caused 
by varying the inside diameter, outside diameter, 
average diameter, or web thickness is given in The 
Interior Ballistics of Rockets . lf 

If only the restriction on internal K is consid- 
ered, one finds that the maximum weight tubular 
grain for a given motor diameter is obtained by 


238 


DESIGN OF ROCKET PROPELLANT CHARGES 


making the axial perforation as small as possible, 
the outside diameter less than 0.6 times the internal 
diameter of the tube, and the length of the grain 
approximately 60 times its diameter. Such grains 
are not practicable, however, for a number of rea- 
sons. Grains with extremely small axial perforations 
cannot be made because the slender “stake” re- 
quired would not withstand the forces encountered 
in extrusion without wandering. The lower limit 
is in the neighborhood of 0.1 times the motor 
internal diameter. Very long grains are not used 
because of the excessive weight of motor tube 
required to encase them and because they are 
obviously ill adapted to stand up without buckling 
under the high longitudinal acceleration forces en- 
countered during firing. 

Empirically, it has been found that the heaviest 
tubular three-ridge grain of JPN ballistite which 
will perform satisfactorily at 140 F statically and 
130 F in the field has the following characteristics: 

Outside diameter 0.83 calibers 

Inside diameter 0.13 calibers 

Web thickness 0.30 calibers 

Length 6.64 calibers 

Here the “caliber” is used as a unit of length equal 
to the internal diameter of the motor tubing. This 
unit is used in nearly all discussions of grain size 
because it enables the results to be expressed in a 
form independent of the actual size of the motor. 
The maximum grain dimensions tabulated above 
have been found to be correct for 2-in. and 3-in. 
calibers, 1 and it appears likely that they would be 
approximately correct for any caliber. Thus for a 
4.625-in. motor like the HVAR, the maximum 
tubular grain would weigh slightly less than 20 lb. 

A very useful method of representing the relation 
between grain shapes and weights is that adopted 
in reference 25 m and in Figure 12. Except for the 
single curve marked “cruciform,” all the data in 
the table are for a single tubular grain inhibited on 
both ends and having no radial holes. If one writes 
down expressions for the volume, burning area, and 


1 The heaviest 2-in. grain which has been used is the Mk 16, 
which has dimensions 1.7x0.28x12.5 in., weighs 1.75 lb, 
and is used in the 2.25-in. subcaliber aircraft rocket. Its 
weight and web thickness are plotted in Figure 12. If this 
grain is scaled up by the factor 1.5 appropriate to a 3.0-in. 
ID motor, it becomes 2.5 x 0.4 x 18.8 in. and weighs 5.9 lb. 
This is only slightly shorter than the longest tubular grain 
which would function in the 3.25-in. AR motor having an 
internal diameter of 3.01 in. 

m Figure 12 is copied from this report except for the curve 
on cruciform grains, which had not previously been published. 


port area of a grain and calculates the possible 
powder volume corresponding to a particular value 
of Ki , it is immediately apparent that all linear 
dimensions are proportional to the internal diam- 
eter of the motor tube and the volume is thus pro- 
portional to its cube. Consequently, one can draw 
a set of curves (dashed lines in Figure 12) giving 
the relations between length, volume, and web 
thickness (expressed in terms of dimensionless pa- 
rameters) of grains having a particular Ki, and 
the curves will apply to all motors and every charge 
whose web thickness is everywhere the same and 
whose burning surface remains constant during 
combustion. The restriction that these grains have 
dimensions which allow them to fit into the motor 
tube limits us to particular portions of the curves 
showing volume (or weight) as a function of web 
thickness, the allowable region depending on the 
type of charge. Thus a tubular grain cannot have 
a diameter larger than that of the motor nor an 
axial perforation smaller than zero, so that, unless 
we are willing to work at a different value of internal 
K, we cannot use a grain of dimensions correspond- 
ing to a point in Figure 12 outside the area bounded 
by the curves “MAX OD” and “ID = 0.” In prac- 
tice, of course, one must remain within a somewhat 
more confined region, and it has been CIT’s prac- 
tice to use outside diameters only between 0.8 and 
0.9 times the inside diameter of the tube and thus 
keep the grains short. It is shown in reference 25 
that all charges consisting of combinations of more 
than one tubular grain have maximum volumes less 
than that obtainable with a single tubular grain. 
In fact, it is easy to see that no other grain shape 
can approach the single tubular grain in possible 
loading density if only geometrical factors are con- 
si dered. 

We have seen that, in practice, the cruciform 
shape gives the highest loading density, and it is of 
interest to show its characteristics on the same 
graph with the tubular grains. One less variable 
parameter is available with cruciform grains than 
with tubular, so that a single curve is obtained in- 
stead of a permissible region of the graph. Plotted 
in Figure 12 is such a curve which assumes (in 
accordance with CIT practice) that the outside 
diameter of the powder is 0.91 times the tubing 
inside diameter (to allow for the inhibitor) and that 
45 per cent of the cylindrical surface and both the 
ends are inhibited. The curve shows that the use 
of cruciform does not allow us to get more powder 


IGNITERS 


239 


in a given length. 11 Its sole advantage is that longer 
grains of this shape can be made to perform satis- 
factorily because (1) the inhibited outer surfaces 
are supported by the motor tube throughout burn- 
ing, (2) grains do not need to be weakened by radial 
holes as do tubular ballistite grains, and (3) the 
inhibitors reduce the surface to volume ratio, de- 
creasing the Ki per pound of propellant. 

Also plotted on the graph are points correspond- 
ing to the heaviest tubular and cruciform grains 


0.33 D 3 for cruciform grains. These values may be 
useful as a rough empirical rule. If lower K r is 
desired, the maximum weight is reduced propor- 
tionately. 

2211 IGNITERS 

The function of an igniter is twofold: to heat the 
propellant grain to ignition temperature and to 



0 0.04 0.08 0.12 0.16 0.20 0.24 0.28 0.32 0.36 0.40 0.44 0.48 0.52 0.56 0.60 

W 

Figure 12. Theoretical maximum volume of cruciform and tubular grains as a function of web thickness 
(see Table 1 for definition of symbols). 


which CIT has found practicable. 0 It is possible to 
conclude from the data that the maximum ballistite 
grain which can be put into a motor tube having an 
internal diameter of D in. without exceeding Kr = 
120 is approximately 0.22 Z) 3 for tubular grains and 


n That the values of X for cruciform and tubular grains are 
not identical results from the fact that the cruciform web 
thickness is not strictly uniform, and a sliver is left after 
burning. 

0 The points were plotted in their proper place with respect 
to web thickness and weight ( w and v ). That the values of OD 
and X read from the graph are not quite correct results from 
the simplifying assumptions made in plotting the curves. 


bring the pressure in the motor up to a point where 
grain will continue to burn satisfactorily. It must 
accomplish these purposes with a short and repro- 
ducible delay p at all temperatures at which the 
rocket is to be used and must not subject the grain 
to excessive forces when it ignites. For efficient heat 
transfer to the grain, it is desirable that the products 
of combustion of the igniter include an appreciable 
amount of solids, since the radiation from gases is 
relatively low. At the same time, however, some 

p Short ignition delays are obviously of special importance 
in aircraft rockets. 


240 


DESIGN OF ROCKET PROPELLANT CHARGES 


gaseous products are necessary to increase the motor 
pressure rapidly to the desired value. 

All CIT rockets are ignited electrically. In a 
typical igniter, a squib, approximately 34 in. in 
diameter and 34 in. long, which consists of a clay 
body with a small depression at one end containing 
an electric bridge wire and a heat-sensitive explosive 
material, is placed in contact with the main ignition 


black powder igniters giving shorter delays at low 
temperatures and lower pressures at high tempera- 
tures. They have not been used in service rockets, 
however, because they are thought to be more 
hazardous than black powder igniters and because 
the magnesium is very subject to surface oxidation 
during storage, with resultant deterioration in per- 
formance. Tests of magnesium-potassium perchlor- 



Figure 13. Igniter types. (A) 2-in. brass case with bakelite closure, (B) 3-in. plastic case, (C) metal case 
for Tiny Tim (230 g capacity), (D) HVAR metal case showing inside with clips for holding squibs and bot- 
tom and top views of assembled igniter. (Note that relative sizes are not accurately shown.) 


charge. Originally a booster charge of finely divided 
black powder or flash powder was placed between 
the squib and the main charge, but it was found to 
be unnecessary in properly made igniters. 

Black powder, usually in the FFFG granulation, 
has been used almost exclusively for the main igni- 
tion charge. Ballistite turnings have some advan- 
tages over black powder for static-firing tests, but 
give long ignition delays. 26 Tests of igniters con- 
taining mixtures of magnesium and potassium per- 
chlorate 27 showed them to be distinctly superior to 


ate mixtures as a squib booster for black powder 
igniters showed a negligible improvement in per- 
formance. 28 

The factors involved in making good igniters are 
discussed in more detail in The Interior Ballistics of 
Rockets. 1 ' 1 For short ignition delays, it is required 
that (1) the firing current be above a certain mini- 
mum (about 2 amperes in CIT igniters) so that the 
bridge wire is heated quickly; (2) the ignition charge 
be tightly compacted; and (3) the igniter case be 
strong enough to remain intact until all parts of the 


IGNITERS 


241 


charge have ignited. In the latter regard, consid- 
erable care is necessary, since a strong igniter case 
will give a high pressure peak at ignition, thus re- 
ducing the safety factor of the motor and contrib- 
uting to high-temperature failures, and it may also 
burst with such violence as to fracture the grain. 
These problems are not serious in fin-stabilized 
rockets where there is usually adequate space at 
the front end of the motor to cushion the shock of 
the igniter’s burst. In spinners, however, where 
length is at a premium, the strength of the igniter 
case is very critical, and one may be forced to accept 
a slight increase in ignition delay in order to prevent 
grain fracture. 

Other desirable igniter characteristics include 
ease of fabrication and loading, ruggedness and 
resistance to vibration, watertightness, and the 
property of fragmenting in such a way as to leave 
no pieces large enough to obstruct the nozzles. The 
types of igniters which have been used in service 
rockets are shown in Figure 13 and discussed briefly 
below. 

The earliest CIT service rockets contained brass 
can igniters with bakelite closures. 29 A drawn brass 
can containing the powder and the squib was 
crimped over a close-fitting bakelite disk which was 
perforated for the squib leads. The crimping opera- 
tion compacted the powder to the desired degree, 
and the igniter was reasonably sturdy. Its disad- 
vantages were frequent squib breakage and poor 
resistance to moisture, and it is now considered 
obsolete except for experimental work. 

Igniter cases of molded plastic have been used 
extensively. 1 *’ 30 - 34 They provide good support for 
the fragile squib by enclosing it in a special com- 
partment and, having threaded closures, allow the 
charge to be very firmly compacted so that their 
resistance to impact and vibration is very good. 
They can withstand complete immersion in water 
for several days. For single-nozzle ground-fired 
motors of 2-in. and 3-in. caliber, they are com- 
pletely satisfactory. For smaller motors and spin- 
ners, however, they cannot be used because the 
cases, in order to be sufficiently strong, must have 
walls approximately 0.1 in. thick with numerous 
reinforcing ribs of greater thickness, so that rela- 
tively large fragments are produced when the case 
breaks up, and these may plug the nozzles. The 
squib compartment is especially bad in this regard 
since it is thick and usually remains intact. In a 
single-nozzle motor, a plugged nozzle means a motor 


burst. In spinners, the primary effect is a decrease 
in accuracy, although bursts may result in extreme 
cases. Finners larger than 1.25 in. in diameter have 
not been made with nozzles small enough to be 
plugged by igniter fragments, but even here the 
fragments may be a disadvantage since they are 
ejected through the nozzles at high velocity and 
may damage the tail surfaces, radiators, etc., of 
aircraft. In order that the case may open up at 
pressures small enough to do no damage to the 
grain, the closure must be made with internal 
threads on the case. 

Igniters with metal cases can be made fully as 
waterproof as those of molded plastic, are even more 
resistant to mechanical stresses, are not affected by 
nitroglycerin (as are some plastics), do not break 



Figure 14. Crimps for metal case igniters: (A) 
standard double crimp, (B) false crimp. 


into large fragments when fired, and are especially 
cheap and simple to make. The cases have been 
made of 0.010-in. tin-plated steel, which is the 
same weight as the material for ordinary tin cans; 
in fact, standard sizes of commercial cans can some- 
times be used. In igniters for large rocket motors, 
it has been the practice to include two squibs wired 
in parallel, thus considerably reducing the number 
of misfires since squib failure is their most important 
cause . 

For finned motors, tin plate igniter cases have 
been made with the standard double crimp (A of 
Figure 14) which is used for commercial cans. This 
crimp is strong and requires considerably pressure 
inside the case before it fails, thus giving the short 
and reproducible ignition delays which are essential 
for aircraft rockets. Although they burst with some 
violence, such igniters do not injure the grain 
because of the cushioning effect of the free volume 
at the front end of the motor. Spinners, having less 


242 


DESIGN OF ROCKET PROPELLANT CHARGES 


than 1 in. between the front end of the grain and 
the base of the head, require an igniter case which 
opens up at much lower pressures, and for these the 
so-called false crimp (B of Figure 14) has been used. 

Usually it is desirable to place the igniter at the 
front end of the motor so that the products of its 




Figure 15. Grid types for single-nozzle motors: 

combustion come in contact with the full length of 
the grain. For spinners, which are necessarily short 
and fat, rear end initiation is possible and desirable 
because the igniter can be put into a space which 
would not otherwise be occupied and need not sub- 
tract from the grain length. Design and tests of a 
toroidal igniter to fit around the grid stool in the 


3.5-in. spinner are discussed in reference 35. It was 
made from plastic, but not being in contact with the 
grain or subject to any compression, its walls could 
be made very thin so that no nozzle-plugging frag- 
ments were produced. The igniter worked success- 
fully, but it was developed too late for service use. 




(A) stool, (B) box, (C) triform, (D) cruciform. 

The size of an igniter charge must be determined 
empirically. Too small an igniter will not give 
reliable ignition at low temperature, and too large 
a one will raise the pressure considerably above the 
equilibrium pressure at high temperatures and thus 
contribute to motor bursts. With the 11.75-in. 
motor, a special igniter problem arose in connection 


GRIDS 


243 


with the shock wave effect on the aircraft structures 
and caused a drastic reduction in the size of the 
igniter charge. (See Section 19.5.2.) 

22 12 DESICCANT BAGS 

Because ballistite is hygroscopic and its burning 
characteristics are dependent on its moisture con- 
tent, the practice of inserting a small bag of silica 
gel in the nozzle exit cone was adopted for the first 
service motor, the ASR, and became standard pro- 
cedure. For small motors the efficacy of such a 
desiccant bag is doubtful at best, and for motors 
where the propellant weight is several pounds the 
bag certainly does no good. In the latter case, the 
moisture capacity of the propellant exceeds that 
of any desiccant bag of practicable size, so that the 
only way to be sure that the propellant has the 
proper moisture content is to load it into the motor 
when its moisture content is correct and then seal 
the motor securely so that it cannot change. Thus, 
if the seals hold, the desiccant is unnecessary, and, 
if the seals spring a leak, the desiccant is not likely 
to be equal to the task of keeping the powder dry. 

With the advent of multiple-nozzle motors, in 
which there was no convenient place to put it, the 
desiccant bag was abandoned. Probably it could 
also be dispensed with on most of the 3.25-in. motors. 

2213 GRIDS 

The purpose of the grid is to support the grain 
and allow free access of the gas to the nozzle. Its 
shape is practically dictated by the shapes of the 
grain itself and of the nozzle, and the principal 
problem in its design is to make it of the right 
thickness so that it will not be too heavy or obstruct 
the gas flow too much and still be strong enough 
at the end of burning to withstand the forces on it 
despite the considerable heating and erosion to 
which it is subject. 

Typical grid shapes which have been used are 
shown in Figure 15. The three-legged (or some- 
times four-legged) stool type and the so-called box 
grid (A and B in Figure 15) have been used for 
tubular grains. The former can be made of cast 
iron or cast steel, but must still be machined on the 
front and rear surfaces and the outer diameter. Al- 
though the stool grid offers better support for the 
grain, the box has usually been preferred because it 


cannot be put in upside down. q Box grids 36 have 
been made by stamping the pieces from steel strip, 
assembling them, and machining to diameter. They 
can also be cast or sintered. The sintering method 
offers the important advantage that no machining 
is necessary, but the pieces usually have low density, 
low strength, and extremely little erosion resistance, 
so that rigid inspection of them is necessary. 

Grids for triform and cruciform grains in 3.25-in. 
motors are shown in C and D of Figure 15. In the 
case of external-burning grains of this type, the grid 
must perform the additional function of keeping the 



Figure 16. End view of Mk 13 grain showing 
wells in end washer to accommodate grid pins. 

grain from rotating and thus closing up the port 
area. The most successful method that has been 
found for assuring this is to have two steel pins pro- 
jecting above the surface of the grid and indexing 
into holes in the grain and its plastic end washer 
(see Figure 16). In cases where it is convenient for 
loading, the grid has been cemented to the end 
washer as an added precaution. The most satis- 
factory production methods for making cruciform 
and triform grids have been steel casting, torch 
cutting from plate, and copper-brazing laminations 
stamped from 11-gauge sheet. 

More complicated grids are required for multiple- 
grain motors such as Tiny Tim. 

<*■ At one time early in World War II, the discovery of one 
ASR motor with an upside-down grid impelled the X-raying 
of several thousand motors. 


Chapter 23 

MOTOR DESIGN 

By C. W. Snyder 


23i INTRODUCTION 

I n this chapter we shall discuss the problems 
encountered in designing the various components 
of a rocket motor and the solutions for them which 
have been used at CIT. Two cautions which apply 
throughout this book should perhaps be specially 
emphasized here. The rocket motors in which we 
have been interested have all been of a special and 
very similar type, namely, those having pressures 
seldom out of the range 1,000 to 2,000 psi at 
ordinary temperatures, burning times in the range 
0.2 to 1.5 seconds, and velocities either subsonic or 
only slightly supersonic. Hence the solutions which 
we have found must not be thought automatically 
to apply to rockets differing too much from these 
specifications. Second, we must ask to be judged, 
in many cases, by our words and not by our deeds 
as embodied in service rockets, since far too often 
important design features were settled by expe- 
diency in the war situation rather than by the ideal 
and, perhaps less frequently, they were settled on 
the basis of preliminary information which did not 
prove finally to be correct. 

23.2 TUBES 

23,2,1 Tubing Dimensions 

The size of a motor tube is ordinarily determined 
by one or both of the following considerations: it 
must fit a particular propellant grain, thus deter- 
mining its length and inside diameter; it must fit a 
particular head, a consideration which, if it exists, 
usually determines the outside diameter, at least 
approximately. The accuracy required on any of 
these three dimensions is never very great. The 
inside diameter must fit the grain, but the accuracy 
with which grains can be made in practice is usually 
less than the commercial tolerance on tubing diam- 
eter, especially when the tubing is made to an ID 
specification, and clearances in the neighborhood of 
% in. on the radius are not objectionable. The 
tubing diameters become particularly critical only 


when peripheral variations in wall thickness cannot 
be tolerated either because of weakening the tube, 
as in the case of an ultrahigh-performance motor 
where the absolute minimum wall thickness is being 
used to save weight, or because of the unbalance 
introduced thereby, as in a low-dispersion spinner. 
In either of these cases, the OD, and perhaps also 
the ID, must be machined since concentricity toler- 
ances, particularly on seamless tubing, are always 
rather large. 


23,2,2 Tubing Material 

For tubing material, nothing other than steel has 
been given serious consideration since alternatives 
which can begin to compete in price and abundance 
do not have the requisite strength and high melting 
point. Seamless tubing is definitely preferred be- 
cause there appears to be no simple and foolproof 
method for detecting a defective weld in a motor 
tube — except possibly by fabricating it into a rocket 
motor and firing it. This difficulty with welded 
tubing was most troublesome with the 3.25-in. AR 
motors, more than half a dozen of which opened up 
at the seam during high-temperature firing, even 
though they had all been hydrostatically tested at 
4,000 psi and the pressure during firing apparently 
did not reach half this value. It seems probable 
that this is to be explained by the more sudden 
application of the pressure during firing, since the 
motors burst before having time to get warm. 


23,2,3 Wall Thickness 

For calculation of the wall thickness and tensile 
strength required, Barlow’s formula is adequate, 
as the wall thickness is always small relative to the 
diameter and great accuracy is not required because 
of the large safety factor which is included. This 
formula is 


t = 


DP 
2 S’ 


( 1 ) 


244 


TUBES 


245 


where t and D are the wall thickness and the outside 
diameter in inches, P and S are the internal pres- 
sure and permissible tensile stress in pounds per 
square inch. It has been the practice to subject 
each length of motor tubing or each completed 
motor to a hydrostatic test at a pressure exceeding 
the maximum normal pressure at the upper service 
temperature limit by a factor 1 .5 for ground rockets 
or 2.0 for aircraft rockets. (With improvements in 
rocket propellants, such large safety factors may no 
longer be justifiable.) The pressure which must be 
used in Barlow’s formula is thus not the motor pres- 
sure but the test pressure. The tensile strength 
to be used is the yield strength rather than the 
ultimate strength, since a motor is not considered 
to have passed the pressure test if it swells by more 
than a specified amount. 

Specification of the test pressure and the motor 
caliber determine, by Barlow’s formula, only the 
product of wall thickness by yield strength. For a 
low-performance motor, because the weight is not 
critical, one usually plans to use ordinary cold-rolled 
steel tubing, for which 50,000 psi is a reasonable 
value of tensile strength, and make the wall thick 
enough to stand the pressure. If high performance is 
required, one usually prefers to use the highest 
grade of heat-treated steel available and make the 
wall as thin as possible to save weight. How far 
one can go in this direction depends upon the heat- 
ing effect. 

23,2,4 Heating Problems 

As the propellant gas flows past any part of the 
rocket it will transfer heat to the surface mainly by 
conduction and convection. Heat will also be trans- 
ferred by radiation, but with ballistite and other 
smokeless propellants the radiative transfer is so 
small a portion of the total heat transfer that it can 
be neglected. The temperature reached by the 
surface depends on the rapidity with which the heat 
received from the gas is distributed by conduction 
throughout the volume of the solid. Ultimately an 
equilibrium would be established in which the rate 
of heat transfer to each unit area of the surface 
would equal the rate of conduction away, but with 
the short burning times characteristic of CIT rock- 
ets, we have to do with transient conditions. The 
theory of heat transfer and conduction is applied 
to rocket motors in The Interior Ballistics of Rock- 


ets , la and in reference 2. We shall consider only the 
results here. 

The time rate' of heat transfer from the gas to the 
rocket’s inner surface is proportional to the dif- 
ference in temperature between the gas and the 
metal and to the heat transfer coefficient h. The 
transfer coefficient is very nearly proportional to 
the “mass velocity” G, defined as the mass of gas 



Figure 1. Tensile strength of metals at various 
temperatures. 


flowing in 1 second through unit area normal to the 
direction of flow. a Thus the heat transfer during 
burning to a unit area is least at the front end of the 
motor, where G is small because gas is practically 
stagnant, and is greatest at the nozzle throat and 
at the nozzle end of the grain, where G is greatest 
because the port area is small and all the gas passes 
by. During burning, the mass velocity at the nozzle 
end of the grain decreases rapidly because of the 
increasing port area, and the rate of heat transfer 

a Actually it depends on the 0.8 power of the mass velocity 
through a proportionality factor which is slightly greater for 
small gas flow channels than for large. 


246 


MOTOR DESIGN 


to the surface decreases in proportion. Hence the 
inside surface of the tube reaches its maximum 
temperature during the first half of the burning time 
and remains nearly constant thereafter, but the 
average temperature of the wall, which is im- 
portant from the standpoint of strength, increases 
steadily throughout burning. 

The variation of tensile strength with temperature 
for typical metals is plotted in Figure 1, taken from 
reference 3. Since the average temperature of a 
motor tube increases steadily during burning, a 
curve of burst strength as a function of time during 
burning would have a very similar shape, with the 
scale depending upon the type of steel, the wall 
thickness, and the mass velocity. If the burning 
time is long enough , the burst strength will eventu- 
ally fall below the motor pressure, and the motor 
will fail. It is thus of considerable importance in 
design to be able to predict the average temperature 
of the critical point of the motor wall at the end of 
burning. By using a value of h which has been 
determined experimentally from a similar rocket, 
this can be done with considerable accuracy, but 
the method is involved and will not be given 
here. 4,5 Typical results of such calculations are 
given in Tables 1 and 2. 

Table 1. Effect of firing temperature on heating of 

11. 75- in. rocket motor wall. 


Wall thickness, 0.280 in. 


Firing temperature (°F) 

-10 

140 

Average reaction pressure (psi) 

960 

1,900 

Reaction time (seconds) 

1.40 

0.70 

Calculated metal temperatures at 



nozzle end of grain at end of re- 



action (°F) 



Inner wall surface 

2000 

2400 

Outer wall surface 

440 

300 

Average 

1040 

1000 

Average temperature rise (°F) 

1050 

860 

Total heat transferred to wall 



(Btu per sq ft) 

1560 

1280 


The total amount of heat transferred to the motor 
wall is slightly greater at low powder temperatures 
than at high because the decrease in the rate of 
heat transfer is more than compensated by the 
greater burning time. At a given powder tempera- 
ture, the total transfer depends very little on the 
thickness of the motor wall. Consequently, as 
illustrated in Table 2, the temperature reached by 
the motor tube is considerably greater for thin- 
walled tubes than for thick, because of the smaller 


Table 2 . Calculated temperature distribution in motor 
wall of 5.0-in. rocket motors at nozzle end of grain.* 


Type of motor 

Mk 2 

Thin-walled 

Thin-walled 

with 

refractory 

Wall thickness (in.) 

0.188 

0.120 

0.120 

Refractory thickness (in.) 



0.010 

Temperatures at end of 
reaction (°F) 

Inner refractory surface 



3250 

Inner metal surface 

2000 

2150 

900 

Outer metal surface 

700 

1550 

500 

Average metal 

1200 

1770 

650 

Total heat transferred to 

wall (Btu per sq ft) 

1050 

1040 

330 


* Assumed properties: 

k 

(Btu/ft • hr • °F) 
Steel 25 

Refractory 0.6 


C 

(Btu/lb • °F) 
0.13 
0.2 


P 

(lb/ft*) 

490 

160 


heat capacity of the thin wall. Whether it is possible 
to achieve a significant saving in weight by using 
high-tensile steel and thin-walled tubes depends 
very markedly on the heating effect. Thus, in the 
example of Table 2, it would probably not be 
possible because at 1770 F no steel would have any 
appreciable strength. 

It should be noted that for a given burning time 
and mass velocity of gas, the absolute thickness of 
the wall, and not the ratio of the wall thickness to 
the diameter of the motor, is the determining factor 
in establishing the temperature. The strength of 
the motor with respect to internal pressure, on the 
other hand, depends upon the ratio of wall thick- 
ness to diameter. Therefore, in small-diameter 
motors of fairly long burning times, the minimum 
wall thickness is generally determined as much by 
the heating effect as by strength requirements so 
that material of unusually high tensile strength 
offers no great advantage. With the larger units 
such as the 11.75-in. motor, a wall thick enough to 
have adequate cold strength is of ample thickness to 
keep the temperature within reasonable limits, and 
considerable weight reduction can be made by using 
high-strength steels. 


Refractory 

The amount of heat which the steel wall must 
absorb can be much reduced by insulating it with a 
thin layer of refractory material. A typical re- 
fractory may have about one-fortieth the heat con- 
ductivity of steel, so that, ideally, the addition of a 
very thin layer of refractory on the inside would 


TUBES 


247 


reduce the wall temperature as much as would a 
considerable increase in wall thickness, with its 
attendant weight increase. Practically, however, 
the low conductivity of the refractory causes its 
inner surface to approach the temperature of the 
gas, and its relatively low tensile strength causes it 
to be eroded away fairly rapidly, so that its effec- 
tiveness does not approach the theoretical value. 
It is possible, nevertheless, to achieve a significant 
saving in weight by the use of refractory and high- 
strength thin- walled tubes, as is shown by tests at 
CIT 2 and by the British experience with the 
RP-3. It was not felt that, in the tactical situations 
for which CIT’s rockets were developed, the in- 
crease in performance attainable by refractory coat- 
ings justified the increased complexity of manu- 
facturing. Hence no very extensive investigation of 
refractories was made . In the future their use may 
be desirable and will change many of the conclusions 
in this book. 

23.2.6 Internal-Burning Grains 

It should be noted that, throughout the discussion 
of the heating effect, it has been assumed that the 
propellant burns on the outer surface so that the gas 
is in contact with the wall. Near the end of World 
War II, as pressure for production slackened and 
more time was available for propellant research, 
experiments with interior-burning grains were begun 
at CIT. At the time of this writing, the continua- 
tion of these experiments at NOTS, Inyokern, in- 
dicates great promise for this design for high-per- 
formance motors where somewhat longer burning 
times are permissible. British research on interior 
burning was already well advanced by the end of 
World War II. The heating effect on such motors b 
is entirely negligible, and aluminum motor tubes 
are feasible. This change also will make a sig- 
nificant difference in the performance attainable 
with rockets of a given caliber. 

232 7 Weldability 

It is frequently desirable to employ welding for 
attaching nozzles, fins, or lugs to motor tubes. 

b High-impulse motors, using internal-burning grains, were 
developed in small sizes at the Allegany Ballistics Laboratory 
[ABL] under Section H, Division 3, NDRC. 6 ’ 7 As of October 
1946, the Hercules Powder Company, operating ABL for the 
Bureau of Ordnance, had developed motors using 100-lb 
internal-burning grains. 


With ordinary mild stee^, this introduces no dif- 
ficulty, but in choosing a high -tensile heat-treated 
steel for a high-performance motor, its weldability 
must be considered. No research on this point was 
done by CIT, since the effect of welding on various 
types of steel is well known to metallurgists. In 
general, the very high-carbon steels undergo a 
marked coarsening of the grain structure and be- 
come brittle, so that even very small welds cannot 
be made without preheating the whole tube. For 
example, with the N-80 oil well casing used for the 
11.75-in. AR C motor, it was found impossible to 
weld on even a row of %-in. studs 9 in. apart, while 
on the 5.0-in. high-velocity aircraft rocket [HVAR] 
motor, using NE 8735, no difficulty was experienced 
with the considerable tack-welding required to 
attach two suspension lugs and four fin lugs. The 
difference in composition between these two steels 
is shown in Table 3. It is easy to be overcautious 
in this regard, since a slight weakening of the motor 
tube in local spots apparently causes no trouble if 
the proper steel alloy is used. 


Table 3. Compositions of steels used in 3.25-in. and 
5.0-in. aircraft rocket motors (NE 8735) and in 11.75-in. 
motor (API N-80 casing). 


Element 

NE 8735 

N-80 

Carbon 

0.33-0.38 

0.40-0.43 

Chromium 

0.40-0.60 

0.08 

Manganese 

0.75-1.00 

1.50 

Molybdenum 

0.20-0.30 

0.16 

Nickel 

0.40-0.70 

0.12 

Silicon 


0.025 

Sulphur 


0.040 

Copper 


0.20 


2328 Threads 

Failures of V Threads 

With the exception of the target rocket which 
used a piston ring closure (see Section 18.5.1), all 
CIT rockets have had threads at the front end of the 
motor tube for attaching to the head, and many 
have been threaded also at the rear to take the 
nozzle. The standard V-shaped thread is not well 
suited to the requirements of rocket motors, where 
strong joints between a relatively thin tube and a 
usually much thicker piece of steel are desired. It 
has been almost universally used, however, be- 

c The 11.75-in. aircraft rocket, usually called “Tiny Tim” 
or simply “Tim” for short, was the last and biggest fin- 
stabilized rocket developed by CIT. 


248 


MOTOR DESIGN 


cause of the much easier availability of dies for this 
shape than for any other and because all machine 
shops are experienced in cutting it. For a given 
thread depth, a V thread is weaker against a straight 
end force than, say, a square thread, but this effect 
is not important in rocket motors since no cases are 
known to the writer in which a motor thread has 
been “stripped” in the ordinary sense. The thread 
difficulties that have arisen (and they were rela- 
tively rare) were caused by the expansion of the 
tube by internal pressure and aggravated, in the 
case of a V thread, by the large angle between the 



sure-tight except in rare instances and some leakage 
of gas occurs, it would be expected from the high 
impedance to gas flow of the interstices between the 
threads that the pressure would drop approximately 
uniformly along the length of the thread engagement 
from full motor pressure at one end to atmospheric 
pressure at the other. This has been confirmed 
qualitatively by a static firing experiment with the 
5.0-in. HVAR motor. The effect of this pressure 
gradient is illustrated, on a very exaggerated scale, 
in Figure 2A. The front end of the motor thread is 
floating in a region of high pressure and hence is not 
expanded, whereas the adjacent portion of the head 
is being expanded by the full motor pressure. This 
expansion allows the full motor pressure to creep 
farther along through the thread and further accen- 
tuate the effect. The result is that only the threads 
at the extreme rear are holding the motor and head 
together and only they will be damaged appreciably 
when the pieces separate. 



B 

Figure 2. Diagrams illustrating probable mecha- 
nism of thread failures on rocket motors at high 
pressure for (A) 3.25-in. Mk 7 motor, (B) 5.0-in. 

Mk 1 motor. Arrows denote approximate pressures 
acting on various surfaces, where P is the total 
motor pressure. 

loaded faces so that a large component of the end 
thrust is transferred into radial pressure. With ex- 
ternal threads on the motor tube, the pressure tends 
to make the threads tighter, and, if the piece into 
which the thin tube screws is relatively thick (as is 
usually the case), no trouble is experienced. With 
3.5-in. heads on the 3.25-in. AR motor, d however, 
where the thicknesses of the two threaded pieces are 
comparable, heads have in several cases been blown 
off by abnormally high pressures in high-tempera- 
ture firing with so little damage to either thread 
that the pieces could be reassembled. 

The probable explanation of this phenomenon is 
as follows. Since the threads are certainly not pres- 

d The combination of a 3.5-in. head with the 3.25-in. motors 
Mk 6 or Mk 7 is designated as 3.5-in. aircraft rocket [AR]. 
See Figure 4 of Chapter 19. 


Special Thread Shapes 

For internal threads on the motor tube, the effect 
is obviously much worse (see Figure 2B) because 
both the internal pressure and the large obliquity 
of the loaded faces tend to expand the motor tube 
but have no effect on the heavy piece screwed into 
it. For this reason, consideration was given to other 
thread shapes for the 5.0-in. motors, but experi- 
ments indicated that they would probably not be 
necessary, and experience has confirmed this fact. 
Only on the 11.75-in. and 14-in. motors, e where no 
advantages from the practical manufacturing stand- 
point were realizable with V threads, was a special 
thread shape adopted. Experience with these two 
motors has indicated that the buttress thread used 
on the latter is probably the optimum thread shape 
both with regard to performance and ease of manu- 
facture. The 7-degree angle of the loaded face is 
small enough to be almost certainly less than the 
angle of repose between steel surfaces, so that the 
end thrust produces no slippage and expansion of 
the tube, and yet it is large enough to provide 
adequate tool clearance and allow the use of thread 
hobs of relatively large diameter. 

It is well known that maximum strength is ob- 
tained when the depth of the thread is one-third 
the thickness of the tube. This rule is useful as a 

e The 14.0-in. aircraft rocket motor was a NOTS project 
initiated in 1945. 


TUBES 


249 


guide, although in most cases it is preferable to use 
a standard thread even though its depth deviates 
somewhat from the optimum. 

Alignment 

Ordinary commercial threads cannot be de- 
pended upon to hold parts in accurate alignment 
because of the relatively large clearances necessary 
to assure interchangeability. To eliminate this dif- 
ficulty, it has been CIT’s practice to use relatively 
loose-fitting threads (No. 2) — obviously desirable 
also from the standpoint of easy assembly under 
adverse field conditions — and to depend for align- 
ment upon screwing solidly against a shoulder. The 
gas malalignment (see Section 24.8) sets a lower 
limit of approximately 1 /io degree below which 
improved alignment of the rocket parts does not 
improve the performance; hence we have made it a 
universal policy that any two rocket parts must 
screw together with a malalignment not exceeding 
this figure. With reasonable care and proper 
machining setups, this accuracy is attainable in 
threading operations without increasing the cost, 
but the methods of specifying and checking it are 
difficult to establish. The specification finally 
adopted as the most satisfactory was that the thread 
seating faces (i.e., the ends of the tube) must be 
parallel to each other and perpendicular to the mean 
axis of the tube within V 20 degree and that a “go” 
thread gauge with a shoulder must seat against the 
tube ends with a gap not to exceed 0.001 in. per in. 
of diameter. f 


2329 Straightness 

The existence of a bow in the tubing has an im- 
portant bearing on the nozzle alignment and hence 
must be controlled. On rocket motors of 3.25-in. 
caliber and smaller, the CIT practice was to bend 
the completed motor g so that the nozzle exit cone 
axis coincided with the center of a “perfect” head 
or of the front end threads within V 20 degree. On 

f That no standard Navy drawings of CIT rockets contain 
this specification is a result of the Bureau of Ordnance rule 
that manufacturing drawings cannot specify gauging methods. 
The usual statement, that ‘‘threads shall align within V 20 de- 
gree” is almost meaningless operationally and has caused con- 
tinual confusion. 

* The apparatus employed for bending tubes is described in 
references 8 and 9. 


the 3.25-in. AR motors, the bend was made at or 
near the front end of the nozzle (which is obviously 
where it belongs), but shorter motors were bent 
approximately at the middle for practical reasons. 

Motors 5 in. in diameter and bigger were not 
practicable to bend, so alignment was secured by 
specifications on the tubes and nozzles and their 
threads separately. Tubing lengths in which the 
bow was excessive were straightened prior to 
machining. 

23.2.10 Reaction with Propellant 

Because corrosion of steel is very rapid in contact 
with smokeless powder, it is necessary to prevent 
the grains from touching the bare motor walls. This 
has been done by painting the inside of the motor 
tube either with standard Navy projectile-cavity 
paint or with clear ethyl cellulose lacquer. The 
lacquer is probably preferable because it gives a 
smoother and harder finish. 

23.2.11 Spinner Motor Tubes 

Only two important factors enter into the design 
of spinner motor tubes which do not appear with 
finners. The first is the requirements of the bourre- 
lets. On the three calibers of spinners which were 
tested by CIT, three types of bourrelet were used. 
Five-inch spinners had the bourrelets on the tube, 
which was machined full length on the outside. On 
the 3.5-in. spinners, the rear bourrelet was the 
nozzle ring, and the front bourrelet was the rear of 
the head, which was slightly larger in diameter than 
the rest of the head. Some experimental 2.25-in. 
spinners were of uniform diameter over the whole 
length, having no bourrelets. 

For barrage or general purpose spinners, any of 
these methods is probably satisfactory. The dif- 
ficulty with uniform diameter rounds is that the 
tubing is never straight. If the outside of the tube is 
machined, the stress relief resulting from the ma- 
chining accentuates the bow in the tubing and 
makes the use of bourrelets necessary for reasonable 
accuracy. 

In the development of the aircraft spinner, there 
was some evidence that a very high degree of accu- 
racy of the bourrelets is necessary (maximum ovality 
not more than 0.002 or 0.003 in.) to obtain minimum 


250 


MOTOR DESIGN 


dispersion. Like all questions of accuracy, how- 
ever, it is difficult to settle, and research into the 
causes of spinner dispersion did not reach the point 
where any general rules could be stated regarding 
permissible tolerances. 

The second problem peculiar to spinners is the 
centrifugal force. Its effect is a peripheral tension 
in the tube which acts like an internal pressure. 
The magnitude of this pressure is easily calculated 
by considering the centrifugal force on unit area of 
the tube and is given in absolute units by 

P = prts\ (2) 

If we take the density p to be 7.3 g per cu cm, 
measure the radius r and thickness t in inches, and 
specify spin velocity s in revolutions per second, 
this is 

P = 0.027 rts 2 (in psi). (3) 

The fastest spin rocket developed by CIT was the 
5.0-in. /14 GASR Model 39A, having a maximum 
spin at 70 F of 309 rps, which gives for the pressure 
equivalent to the centrifugal force approximately 
840 psi. Apparently the effect of centrifugal force 
on the motor tube becomes important only at ex- 
tremely high spins, but its effect on powder breakup 
does cause motor bursts as has already been noted. 

233 NOZZLES 

The functions of a rocket nozzle from the view- 
point of interior ballistics has been discussed in 
Chapter 21, where it was shown that the important 
characteristics of a nozzle are its throat diameter, 
which determines the equilibrium pressure of the 
motor, and its expansion ratio, which determines 
the amount of additional thrust which can be wrung 
out of the gases during their expansion. This addi- 
tional thrust, expressed quantitatively by the nozzle 
coefficient Ch is determined in practical cases by the 
expansion ratio, since the divergent angle of the 
exit cone is never made so large that its effect is 
appreciable. Obviously, since the gas in the throat 
is moving with the velocity of sound, a nozzle with 
a 45-degree half-angle of divergence would have a 
very low nozzle coefficient regardless of its ex- 
pansion ratio, since the gas could not expand 
rapidly enough to touch the exit cone at any point. 
Half-angles from 6 to 15 degrees have been used, 
and little is known of the behavior of more rapidly 
diverging nozzles, although it is probable that they 


would give decreasing accuracy as well as decreasing 
thrust. 

The ideal interior contour of a nozzle is deter- 
mined by the desire for maximum accuracy and 
minimum erosion. Both these considerations favor 
very long nozzles with gradually tapering entrances 
and exits. It has been repeatedly demonstrated 
that whenever the gas is required to change its 
direction abruptly, local erosion is severe. On 
accuracy, the evidence is less clear-cut, but it ap- 
pears that the exact contour of a nozzle is unim- 
portant provided that (1) it possesses axial sym- 
metry and (2) that the flow of gas delivered to it is 
uniform. 10 In practice, however, the gas flow to the 
nozzle is not uniform because of the complicated 
shapes of grains and grids; hence longer nozzles 
give better accuracy because they have more time 
available for straightening out this nonuniform 
flow. 11,12 

Considerations of space, weight, and ease of fab- 
rication dictate that nozzles are always made short 
and with simple contours. Thus the exit portions 
are always conical, and the entrance is a combina- 
tion of straight lines and circular arcs. It has never 
been possible to obtain a clear correlation between 
dispersion and any characteristic of the entrance 
portion of the nozzle other than its alignment . The 
varied shapes which exist have resulted from con- 
siderations of manufacturing methods, necessity for 
fitting grids, and esthetics. 

23 3 1 Nozzle Types 

It is difficult to lay down any very useful general 
rules for deciding which type of nozzle is preferable 
for a particular rocket. In CIT’s case, the choice 
tended to be influenced greatly by the type of 
machine tools that were available to us at the time , 
since the project was doing both design and pro- 
duction. For fin-stabilized rockets, the basic choice 
is between single nozzles and multinozzles. Single 
nozzles are obviously the choice for small motors 
(2.25-in. and smaller) because they are simpler and 
cheaper to manufacture and because, with multiple 
nozzles, each nozzle would be so tiny that its erosion 
would be large. In the large calibers (5.0-in. and 
larger) the advantage in ease of manufacture prob- 
ably lies with the multinozzle and three other ad- 
vantages become important: 

1 . The possibility of having a central nozzle with 
a blowout disk, thus increasing the safety at high 


NOZZLES 


251 


temperatures and greatly extending the usable tem- 
perature range ; h 

2. A considerable saving in length and perhaps a 
slight saving also in weight, although the latter is 
uncertain since no single-nozzle large motors have 
been made; 

3. A decrease in dispersion resulting from the 
averaging out of “gas malalignment” between the 
various nozzles. 

In the intermediate sizes (3.25-in.), the choice is 
difficult. That only single nozzles have been used is 
an indication not of their superiority but of the fact 
that the advantages of multinozzles became apparent 
gradually during the existence of the project. For 
the 3.25-in. AR motor, for example, the abandon- 
ment of the single-nozzle design in favor of multi- 
nozzle was recommended to the Bureau of Ordnance 
by CIT early in 1945, and a thorough investigation 
would probably reveal that some of the other rock- 
ets of this caliber could be improved by the change. 
The best argument for multinozzles — the blowout 
disk — is, however, less cogent for low-performance 
and nonaircraft rockets. 

23 3,2 Single Nozzles 

The simplest way to make a nozzle is to shape the 
rear end of the motor tube into the proper contour 
(see Figure 4A). Such “integral” nozzles have been 
used extensively by the Army, whose rockets have 
relatively thicker walls than CIT’s, and were used 
on several early rockets. In some instances, nozzle 
and tube were made separately and butt-welded 
instead of being formed from a single piece. On the 
target rockets they were satisfactory because the 
accuracy was almost completely controlled by the 
fins, but on the CWR (see Section 18.4) they were 
abandoned for accuracy reasons. It was never 
possible to manufacture them without appreciable 
variations in thickness around the nozzle throat. 
Hydrostatic pressure tests 13 and experiments with 
the yaw machine 12 showed that these variations 
caused the exit cone of the nozzle to deflect under 
the pressure of the firing, thus changing the axis of 
thrust. If a really good fabrication method were 
available, integral nozzles would have important 
advantages in saving weight and eliminating several 
manufacturing operations, but it seems clear that 

h Several gadgets for achieving the same result with a single 
nozzle were tried but showed little promise. 


the ordinary methods of swaging and spinning can- 
not make nozzles of sufficient accuracy, at least in 
the range of wall thickness which has been in- 
vestigated. 

The inaccuracy obtained with integral formed 
nozzles is largely eliminated if both ends of the 
nozzle are held firmly by a piece of tubing so that 
the exit cone cannot deflect appreciably. Hence 
separate formed nozzles 1 inserted into the motor 
tube have been used successfully on the majority of 
CIT rockets. They have been made rapidly and 
cheaply by a number of techniques 14,15 with suffi- 
cient accuracy to be acceptable, although again the 
chief difficulty with them is accuracy. 

Nozzles machined from bar stock were used, on 
the antisubmarine rocket [ASR] and barrage rocket 
[BR], for example, before acceptable techniques for 
forming nozzles had been developed. Functionally 
they are preferable to any other, since they can be 
made as accurately as desired, but they cannot 
compete in mass production with the formed nozzle 
except in small sizes where screw machines are 
readily available. Thus in CIT production the 
formed nozzle for the 2.25-in. SCAR j cost 75 cents 
to make and 25 cents to braze into the tube. The 
nozzle for the BR was very similar, but, machined 
from bar stock, it cost more than twice as much. 

Attachment of Single Nozzles 

For attaching machined nozzles to motor tubes, 
two methods have been used in quantity production, 
as shown in Figure 3. The use of threads is prob- 
ably not ideal because of the objection to internal 
threads on the motor tube discussed under Special 
Thread Shapes, in Section 23.2.8, and because it is 
difficult to be certain that the threaded joint is 
moisture-tight. Unless care is taken to tighten the 
nozzle firmly, it may move slightly when the pres- 
sure comes on the tube, thus introducing a malalign- 
ment. Nevertheless, threaded nozzles were used 
extensively on low-performance motors and were 
satisfactory. The specification of thread alignment 
with the seating face naturally applies to the nozzle 
threads as well as to the tube. Threads cannot be 

* In early CIT reports this type of nozzle is often called a 
re-entrant nozzle. 

j The latest official designation of the forward-firing practice 
rounds is “2. 25-in. Forward-Firing Aircraft Rockets (Target).” 
In most of the literature they are known as subcaliber aircraft 
rockets [SCAR] . Three variations are distinguished by Mark 
numbers. See Figure 7 of Chapter 19. 




252 


MOTOR DESIGN 


used with formed nozzles because the thin wall will 
not accommodate the necessary seating shoulder. 

For 1.25-in. motors, the swaging method shown 
in Figure 3B is probably the best solution. The 
joint is rigid when properly made and is well adapted 
to quantity production. 



A ASB AND BR 


B ALL 1.25 -IN. 

MOTORS 

Figure 3. Methods of attaching machined noz- 
zles. 

Copper brazing or silver soldering has been used 
for most formed nozzles and for a few machined 
nozzles. A smooth joint is formed and, particularly 
with induction heating, the rate of production is 
good, and the damage to the motor tube by the 
heat is negligible because the critical part of the tube 
(just ahead of the nozzle) does not get very hot. 
With the relatively thin formed nozzles it is desir- 
able that both ends fit snugly into the motor tube 
since otherwise one runs into the same warpage dif- 


ficulty as with the integral formed nozzle, although 
on a reduced scale since here it would be the entrance 
of the nozzle rather than its exit cone which would 
be shifted by the warpage. To avoid having to press 
the nozzle in for its full length, a procedure which 
usually results in galling the inside of the tube and 
rolling up metal ahead of the nozzle so that the grid 
does not seat properly, three alternatives have been 
used (see Figure 4) . k 

1. On the 2.25-in. SCAR, the rear end of the tube 
was machined internally for the length of the nozzle 
to a diameter nominally equal to the nozzle external 
diameter, so that clearance or interference up to 
0.004 in. was possible in the most adverse cases. 
The fact that the front end of the nozzle could have 
a few thousandths of an inch freedom was accepted 
in the interest of easier production, since the nozzles 
were relatively thick in proportion to their diameters 
and the accuracy of a practice round was not of 
prime importance. Heavy press fits were eliminated 
by selective assembly when necessary. 

2. On the 3.25-in. AR Motor Mk 7, a bead was 
rolled or pressed into the tube so that the nozzle 
would drop into the tube loosely from the rear and 
be tight for the last one-quarter inch approximately. 
In order to meet the two requirements that the rear 
end be a press fit and that there be a small clearance 
for the silver solder, a 0.002-in. step was machined 
on the rear contacting surface of the nozzle. This 
method of attachment was evolved after consider- 
able experience with others and is believed to be 
the best. 

3. On the VAR series * 1 (3.25-in. Motor Mk 1 
et al.) , the nozzles were made as shown in Figure 4D 
because leaving the tube with its full 3.25-in. diam- 
eter at the rear allowed so little airflow through the 
7.2-in. tail that the stability of the rockets would 
have been unduly low. The same design was 
adopted for the first AR motor (3.25-in. Mk 6) 
in the interest of standardization but soon abandoned 
because it has little to recommend it. The com- 
plicated shape was much more difficult to make than 
the bead in the Mk 7 motor, and the reduction of 
the tube diameter ahead of the nozzle was undesir- 

k Several other possibilities were tried on the BR but were 
abandoned because of increased dispersion. They are illus- 
trated in reference 16. 

1 The series now designated “7.2-in. retro rockets,” designed 
for firing backward from aircraft, has more frequently been 
called vertical antisubmarine rockets [VAR]. Velocities of 175, 
200, 210, 310, and 400 fps are obtained with 3.25-in. motors of 
different length but identical design. See Figure 2 of Chap- 
ter 19. 



NOZZLES 


253 






D VAR AND 
3.25-IN. AR Mk 6 

Figure 4. Methods of attaching formed nozzles. 

able because it increased the internal K of the 
motor. On low-performance motors the change in 
internal K was not critical, but on the aircraft 
rocket motor it gave an easily measurable reduction 
in the upper temperature limit. A further disad- 
vantage, which again is most significant for high- 
performance motors because of their large nozzle 


/ 

throat, is the reduction in nozzle expansion ratio 
entailed by the swaging down of the tube . 

Because the silver solder joints were usually the 
weakest point of the motor, it was standard practice 
to give them a thrust test with a force corresponding 
to the product of the internal cross-sectional area of 
the tube by the maximum expected motor pressure 
with an appropriate safety factor. A considerably 
stronger joint can be made by arc welding, as was 
done on the VAR/s and some others, but this tech- 
nique is not favored because it leaves a rough exit 
circle. In addition to the obvious objection of the 
necessity for cleaning up the weld, the roughness 
has a more subtle fault. Since the gas is discharged 
from the nozzle at a pressure above atmospheric, it 
exerts a radial pressure on the nozzle exit cone, and, 
if the cone is slightly longer on one side than the 
other, there will be a net side force which is small in 
magnitude but large in effect because of its very long 
lever arm relative to the center of mass. Tests with 
the 3.25-in. AR 17 indicated that the deflection so 
introduced was of the order of 2 mils per 0.01 sq in. 
of unbalanced area. 

Considerably thinner stock can be used for form- 
ing nozzles if the motor pressure is given access to 
the annular space between the nozzle and the tube. 
On the British RP-3, the annular space is sealed 
from the inside of the motor by an obturator cup 
because it is open to the outside through the fin 
slots. The CIT practice, on the other hand, has 
been to provide ports between the annular space 
and the inside. If this is not done, the combination 
of the pressure gradient between the motor and the 
annular space and the setback force of the grain will 
collapse a thin nozzle at the throat. The holes arc 
placed so that any lubricating oil or cleaning com- 
pound which might be trapped in the annular space 
will drain out when the motor is stood on the front 
end, since otherwise it might seep out after the 
rocket is loaded and react with the propellant. 

23.3.3 Multiple Nozzles 

The first problem facing the designer of a multi- 
nozzle rocket is the number of nozzles to use. For 
fin-stabilized rounds, where malalignment is im- 
portant, the choice is considerably narrowed by the 
rule that the nozzle arrangement should have the 
same symmetry as the grain so that the amount of 
gas flowing through different nozzles is equal or at 


254 


MOTOR DESIGN 


least symmetrically arranged. For example, a motor 
containing a cruciform grain should have four or 
eight nozzles or a multiple thereof, whereas six 
nozzles are appropriate for a triform grain. That 
this rule is necessary is based on good logic and poor 
experimental evidence, but it has been followed be- 
cause it turned out to be convenient to do so. The 
experimental evidence consists of (1) a firing of six 
rounds of four-nozzle CWR’s with three-ridge grains 
which flew wildly for reasons unknown 18 and (2) 
the fact that the nozzles of the 5.0-in. spinners 
which are shielded by the legs of the grain do show 
less erosion in static firing than those opposite the 
openings. 

Consideration of nozzle erosion is important, 
since its effect is much greater on many small 
nozzles than on a few larger nozzles having the same 
total throat area because of the greater exposed 
surface of the smaller nozzles. The change in total 
nozzle area is roughly inversely proportional to the 
nozzle radius, so that, unless it is possible to adjust 
the progressiveness of the grain to compensate for 
the increased nozzle area (as was done on the cruci- 
form charges for 5.0-in. spinners), one will not get 
good burning curves if the nozzle radius is too small. 

With these two factors in mind, one usually 
chooses the number of nozzles primarily on the 
basis of the space available in the nozzle plate. 
There is probably an optimum number from the 
viewpoint of manufacturing cost, since the lower 
unit cost of making a small hole is balanced by the 
larger number- of them required, but this is not a 
very critical criterion. 

Multinozzles can either be machined directly in a 
nozzle plate or made individually and inserted into a 
relatively thin plate. The former “integral” type 
has been used in finners and the insert type in 
spinners because of the disparity between the 
amounts of propellant in the two types. A glance 
at the nozzle plate of an HVAR or a Tim will show 
that so much of the area is taken up with nozzle 
that if one is to have an adequate expansion ratio 
(approximately 4 is usually considered desirable), 
there would be almost no metal between nozzles of 
the insert type, and the plate would not withstand 
the motor pressure. If one were to make a low- 
performance finner with a propellant charge com- 
parable to those which, because of the length 
limitation, are used in spinners, he might choose 
the insert-type nozzle plate. It has been used ex- 
clusively on spinners primarily because of its con- 


siderably smaller weight — approximately 4 lb for 
the 5.0-in. spinner compared to 7.5 lb for the HVAR. 

In the matter of cost, the advantage lies with the 
insert nozzles because a slip in machining one nozzle 
hole does not result in scrapping the whole assem- 
bly. Thus in CIT production of over 100,000 
motors, the one-piece nozzle plate (with its skirt or 
ring) for the HVAR cost $11.87. Despite its much 
greater complexity, the nozzle assembly for an 
eight-nozzle spinner could have been made for less 
than $8.50. 

The individual insert nozzles have been made as 
simple as possible with a cylindrical outer surface 
in. order to keep the cost down. CIT purchased 
5.0-in. spinner nozzles at 10.3 cents each. Putting 
a shoulder on them to keep them from being blown 
out by the motor pressure requires a considerable 
increase in machining cost. Copper brazing was 
universally used for holding the nozzles in the plate, 
although silver solder would be equally good, and 
other suggested methods (such as pinning) appear to 
have no functional disadvantage provided that the 
nozzles are not loose in their holes. 


Nozzle Tolerances 

Since nozzles are difficult to manufacture because 
of their complicated shape and this difficulty in- 
creases greatly as the specifications and tolerances 
are made more stringent, it would be extremely 
useful to be able to define precisely the limits within 
which inaccuracies in fabrication will not noticeably 
affect performance. This is never even approx- 
imately possible in practice because in any border- 
line case it is the dispersion that is in question, and 
dispersion is extremely difficult to measure pre- 
cisely. It is influenced by such a diversity of factors 
difficult to control that, unless the factor being 
considered has a very large effect (as is seldom the 
case), one can seldom say with certainty whether 
the difference in dispersion between two sets of 
field firings was the result of the factor in question 
or not. It is, of course, also true that no borderline 
between good and bad nozzles exists, but all grada- 
tions between best and worst appear. In setting 
standards of acceptance for nozzles, one is thus 
continually required to make arbitrary decisions 
with little or no assistance from the experimental 
facts. A few general principles are available to guide 
the decision, and these are listed in the following 


NOZZLES 


255 


paragraphs. But beyond these, the best that can 
be done is to assume that the ideal nozzle is per- 
fectly smooth and perfectly symmetrical in all de- 
tails and to reject on principle any manufacturing 
method which gives nozzles differing more from the 
ideal than those made by another method. Thus 
hot spinning was abandoned by CIT when other 
forming techniques became available which gave 
smoother interior surfaces, even though the effect 
of smoothness was not very firmly established 
experimentally. 

The throat diameter of a single-nozzle motor 
affects only the operating pressure, but its dimen- 
sion is not very critical because the variation in 
surface area among different grains is usually about 
=hl per cent. A variation of the same amount in 
nozzle throat area corresponds to such a large vari- 
ation in diameter that tighter tolerances have been 
specified on the drawings in order not to encourage 
sloppy workmanship. On multinozzle motors, uni- 
formity of nozzle diameter is required to keep down 
the malalignment. 

The thickness of a nozzle must be great enough at 
every point to withstand the setback force of the 
grain (and also the pressure differential in case the 
nozzle is not vented), but the uniformity of thick- 
ness is important to guarantee that it does not dis- 
tort unsymmetrically when the pressure and heat 
are applied and thus introduce dispersion. 

On machined nozzles it is not feasible to blend the 
entrance and exit cones into a smooth curve, and a 
short cylindrical surface is left at the throat. The 
length of this flat does not appear to influence dis- 
persion if it is small compared to the throat diam- 
eter, but sharp angular transitions between it and 
the conical portions have been avoided lest there be 
a tendency for the gas to pull away from the surface. 
The latter consideration may not be significant be- 
cause a sharp angle would erode away very quickly. 

The surface smoothness is unimportant within 
rather wide limits. Certainly nothing is to be 
gained by honing or polishing the interior of a 
nozzle to a better finish than that of ordinary cold- 
rolled steel (about 100 microinches) 19a and a con- 
siderably rougher finish would probably be satis- 
factory except for the fact that it has not been 
possible to devise a gauge for checking the direction 
of the axis of a rough nozzle. Nothing can be 
learned by firing rough nozzles, since the direction 
of their alignment is not accurately known. Gouges 
or ridges or other imperfections are to be avoided if 


they are unsymmetrical abound the periphery, espe- 
cially if they are in the throat or exit cone. The 
entrance cone appears to have no effect on accuracy 
unless it is displaced or cocked at a considerable 
angle with respect to the throat and exit cone. The 
effective axis of the nozzle is almost exclusively 
determined by the axis of the throat and exit cone. 

Ovality of the throat or exit cone is undesirable 
for the same reason as roughness — the alignment- 
checking mandrel will not determine the actual 
effective axis of the nozzle, and, if this uncertainty 
is much greater than Vio degree on the average, an 
increase in dispersion will result. 

For multiple-nozzle plates on finners, we have the 
additional requirement that the average alignment 
of the nozzles must be perpendicular to the thread 
seating face within the usual V 20 degree, since on 
such large motors it is not practicable to bend the 
tube to bring the nozzle axis into coincidence with 
the center of mass. The alignment of any particular 
nozzle can be allowed to vary by several times this 
amount. A similar requirement is necessary for 
spinners, although here the tolerance depends on 
the stability factor. That the effect is significant in 
practice despite the averaging of the malalignment 
by the rotation was shown by a test on the 5.0-in. 
HCSR Model 134, m in which cocking the nozzle 
plate degree deflected the rocket 14 mils from its 
trajectory for zero malalignment. To guard against 
such a consistent error in cant angle for several 
nozzles in one plate, a fairly close tolerance on cant 
angle was specified. 

23 3,5 Flash Suppression 

The elimination of the luminosity of the rocket 
jet is desirable in some applications for concealment 
and is particularly important for forward-firing air- 
craft rockets, where the flash may temporarily blind 
the pilot during night combat. It was found that 
single-nozzle rockets having small nozzle expansion 
ratios gave very luminous trails during the whole 
of burning, the brightness being greater at higher 
temperatures. A nozzle with a large expansion ratio 
apparently cools the gas below the flash point before 
allowing it to mix with the air, so that the trail is 
invisible except for an instant at ignition and again 
when the grain collapses at the end of burning. 

m In the standard CIT designation, HC denotes the head 
type (high-capacity) and SR denotes spin-stabilized rocket. 


256 


MOTOR DESIGN 


With the divergence angle usually used (6 to 9 
degrees half-angle), the minimum expansion ratio 
for flashless performance at all temperatures is close 
to 4.0, but it appears to be larger for nozzles with 
much larger divergence angles. 19,20 

For multinozzle rockets, the situation is probably 
little different, although no comprehensive investiga- 
tion of the effect of expansion ratio on flash has been 
made. The HVAR, having an expansion of 4.0, was 
rendered flashless by adding a “nozzle skirt,” a 
piece of 5-in. tubing which extended 1 in. behind 
the rear face of the nozzle plate. A multinozzle 
with an expansion ratio of 5 will give flashless per- 
formance as a nonrotating rocket but not as a spin- 
ner. 21 No way is known to eliminate the flash of 
spinners. 

2336 Erosion 

We have already seen that the rate of heat trans- 
fer per unit area from the propellant gas to the 
rocket metal parts is almost proportional to the 
“weight velocity” and hence is largest at the rear 
end of the grain and in the nozzle throat. The 
decrease in tensile strength with high temperature 
does not ordinarily cause collapse of the nozzle 
because the throat is of small diameter and has a 
relatively thick wall, but it does cause erosion which 
increases the nozzle area during burning at a rate 
dependent primarily on 

1. The material of which the nozzle is made. 

2. The temperature of the propellant gas. 

3. The motor pressure. 

4. The burning time. 

5. The shape and size of the nozzle. 

Despite the large discrepancy of approximately 
2700 F between the melting point of steel and the 
temperature of the gas, it does not appear that any 
appreciable melting occurs, since, long before the 
melting point is reached, the strength of the steel 
becomes insufficient to withstand the high stresses 
imposed by the flowing gas, and plastic flow occurs 
in the metal. 22 That this explanation of erosion is 
correct is indicated by the photomicrographs of 
Figure 5 which show a typical case of relatively 
severe erosion such as is encountered with the small- 
diameter spinner nozzles shown in Figure 6. Figure 
5 shows a section near the nozzle throat. The grain 
structure of the metal in zone A, next to the surface, 
shows that it has been heated above the Ac 3 point 



Figure 5. Panorama of inside area of cold-rolled 
steel nozzle. Portion of region near nozzle throat 
shows the changes in grain structure due to the 
temperature gradient. Zone A shows structure 
corresponding to Ac 3 (1560 F). Zone B corre- 
sponds to Aci (1365 F). Zone C did not reach the 
Aci temperature. 


NOZZLES 


257 



UNFIRED 


steel which had not been heated to the Aci point, 
and a superimposed layer of steel which had flowed 
from the throat', having been above the AC3 point 
but not melted. The temperature distribution in 
the nozzle is plotted in Figure 8. It should be noted 
that the temperature is high only in a very thin 
surface layer. 


Figure 7. Photomicrograph (132 x) of section of 
exit cone of steel nozzle. Lower part is cold-rolled 
steel nozzle proper which has not reached the Aci 
point. Upper part is metal which has plastically 
flowed from nozzle throat. 

This analysis of nozzle erosion shows that the 
melting point of a nozzle material is of importance 
only indirectly in that the tensile strength tends to 
be low near the melting point. For a heat-resistant 
nozzle, the important factor is high tensile strength 
in the neighborhood of 2000 F. Two types of heat- 
resisting materials have been suggested — the ceram- 
ics and the high-melting-point metals. No ceramic 
that has been tried even approaches the requisite 
strength. Even with a powder weight of only 1 .5 lb 
as used in the ASR, the ceramics cracked and eroded 
so severely as to reduce the final pressure to less 
than half its normal value, whereas ordinary steel 


(1560 F); that in zone B indicates heating above the 
Aci point (1365 F) but below the Ac 3 point; and that 
in zone C has not been altered, so that the tempera- 
ture must have remained below the Aci point. The 


130 


130 


Figure 6. Typical erosion of small steel nozzles. 


maximum temperature reached in zone A cannot be 
determined metallurgically , but extrapolation would 
indicate a value in the neighborhood of 1800 F. The 
photomicrograph of a section from the same nozzle 
in the exit cone (Figure 7) shows the unaffected 


258 


MOTOR DESIGN 


nozzles on this rocket can be fired several times. 
Molybdenum and tungsten nozzles show virtually 
no erosion at all. Tungsten carbide, which can 
easily be cast into the proper shape, also works 
well, but because of its brittleness it must be properly 
supported. Thus a nozzle throat insert of tungsten 



0 0.10 0.20 
DISTANCE FROM NOZZLE THROAT IN INCHES 
h = 0.15 cal/cm 2 SEC °C 

Figure 8. Temperature distribution in steel noz- 
zles; h = 0.15 cal/cm2*sec*°C. 


which has actually been used may be called the 
heat-absorbing type, which depends upon its 
ability to cool the surface by conducting heat away 
from it faster than the gas can supply it. The most 
important property of such a nozzle material is its 
thermal conductivity. Thus under conditions where 
the inner surface of a cold-rolled steel nozzle would 
reach a temperature of 2040 F and erode away con- 
siderably because its tensile strength becomes effec- 
tively zero about 400 degrees below this, a copper 
nozzle would not get above 950 F and would show 
very little erosion since it would still have some 
strength at that temperature. The theory of heat- 
absorbing nozzles is discussed in reference 23, from 
which Table 4 is taken. Experimental data in the 
last column of this table are taken from tests at 
130 F with the insert nozzles of the 3.5-in. spinner 
where erosion is especially severe because of the 
small throat diameter (0.289 in.). The results of 
these experiments were in complete agreement with 
the theory. Thus various types of high-speed tool 
steel were all found to be inferior to cold-rolled steel 
because their low conductivity more than counter- 
balanced their greater strength. In particular, 
Stellite and Hastelloy, special alloys which maintain 
a high tensile strength even at red heat, gave the 
highest erosion of all the metals tested, the surface 


Table 4. Characteristics of nozzle materials. 


Predicted 


Thermal surface Percentage 


Metal or alloy 

conductivity 
at 1600 F 
(cal/cm- sec- °C) 

Density 

(g/cm 3 ) 

Specific 

heat 

(cal/ g- °C) 

Thermal 
capacity 
(cal/cm 3 - °C) 

Tensile 

strength 

(psi) 

temperature 
h = .22 

0 = .45 
(°F) 

Melting 

point 

(°F) 

Quality as 
nozzle 
material 

erosion 
in actual 
testing 

Hastelloy 

0.03 

8.94 

0.092 

0.92 


2700 

2350 

Very poor 

65 

Stellite 

0.035 

8.38 

0.10 

0.84 


2650 

2370 

Very poor 

59 

Inconel 

0.036 

8.51 

0.109 

0.93 


2700 

2540 

Very poor 

46 

Stainless steel 

0.039 

8.0 




2650 

2700 

Very poor 


Monel K 

0.062 

8.5 

0.127 

1.06 

20,000 at 1600 F 

2200 

2400 

Very poor 


Cr steel 

0.07 

7.74 

0.11 

0.85 

0 at 1800 F 

2100 

2700 

Poor 

45 

Cold-rolled steel 

0.0875 

7.8 

0.168 at 1600 F 

1.31 

0 at 1600 F 

2040 

2600 

Poor 

40 

Tantalum 

0.130 

16.6 

0.036 

0.60 


1750 

5162 

Excellent 


Iron 

0.19 

7.8 

0.162 at 1800 F 

1.26 

0 at 1800 F 

1750 

2795 

Fair 


Molybdenum 

0.346 

10.2 

0.075 

0.78 


1250 

4748 

Excellent 

none 

Beryllium 

0.385 

1.8 

0.505 

0.94 


1240 

2462 

Good 


Chromium 

0.65 

6.9 

0.187 

1.29 


1070 

2939 

Good 


Aluminum 

0.66 

2.7 

0.277 

0.75 

0 at 600 F 

1050 

1218 

Very poor 


Copper 

0.858 

8.9 

0.126 

1.12 

0 at 1000 F 

950 

1981 

Fair 


Silver 

0.97 

10.5 

0.076 

0.80 

0 at 1100 F 

900 

1760 

Fair 



carbide with a cylindrical outer surface cracked 
severely when fired, whereas with a conical surface 
no cracking occurred. None of these heat-resisting 
nozzles have been used because there was not suffi- 
cient need for them to justify the extra cost. 

In contrast to the heat-resisting nozzles, the type 


temperature apparently actually reaching the melt- 
ing point. In Figure 9 is shown a comparison of the 
Stellite nozzle with one of chromium-plated copper. 
The latter works very satisfactorily because the 
copper has an extremely high conductivity (nearly 
ten times that of steel) while the chromium, although 


NOZZLES 


259 


having a somewhat lower conductivity than copper, 
contributes its high melting point and hardness. 

Among the various low-carbon free-machining 
steels which one would naturally choose for ma- 
chining a complicated piece like an integral multi- 
nozzle, there is little difference in erosion charac- 
teristics, but any of them is significantly better than 


Despite the shorter burning time, erosion is 
greatest at high temperatures because of two 
effects: (1) the higher weight velocity increases the 
coefficient of heat transfer from the gas to the 
nozzle wall, and (2) the higher motor pressures 
cause plastic flow to occur at lower nozzle tempera- 
ture. The variation of erosion with powder tem- 



FiGURE 9. Extremes of good and bad nozzle erosion under identical conditions. Left: chromium-plated cop- 
per. Right: Stellite. Initially the nozzles had identical inside contours. 


SAE 1020, apparently because of higher manganese 
content. 24,25 A few sintered and cast nozzles which 
have been tried have all eroded seriously. No really 
comprehensive study of nozzle materials was at- 
tempted by the project because, at the short burn- 
ing times in use, the problem was not of sufficient 
urgency to warrant it. Such surveys have been 
made by groups interested in liquid-fuel rockets and 
jet engines. 26 


perature for the case of the 3.5-in. spinner 27 is 
shown in Figure 10. 

Since erosion is most severe at points where the 
gas is forced to change its direction rapidly, some 
improvement can sometimes be obtained by careful 
attention to the contours. Thus longer entrance 
cones reduce nozzle throat erosion, but this fact is 
of little importance because in practice one prefers 
to have entrance cones as short as permissible. 


260 


MOTOR DESIGN 



WEIGHT FRACTION BURNED 


Figure 10. Increase in area of small steel nozzles 
from erosion during burning at various propellant 
temperatures. 

23 3 7 Blowout Disks 

For large rockets or for those which are used in 
situations where a motor burst would involve 
exceptional hazards (e.g., aircraft rockets), it is 
desirable to include an extra nozzle in the center of 
the plate and close it with a blowout disk which is 
ejected if the motor pressure exceeds a particular 
value. This device is made necessary by the rela- 
tively small strength and large temperature coeffi- 
cient of the present powder, and as rocket propel- 
lants are improved, its use will become less neces- 
sary. It allows one to combine the characteristics 
of two different rockets in one jacket. With the 
blowout disk closing the central nozzle, the nozzle K 
is high, say 210 to 220 for ballistite, so that the 
motor operates at relatively high pressure and short 
burning time, having its range of useful temperature 
displaced below that usually designed into rockets. 
With the central nozzle open, the situation is 
reversed so that high-temperature performance is 
increased at the expense of low-temperature per- 
formance. If increasing the useful temperature 
range were the only consideration, the blowout disk 
would be designed to be ejected at approximately 
the temperature midway between the two extremes 
desired. In practice this has not been done because 
in the vicinity of the blowout pressure it is im- 
possible to predict whether the disk on a particular 
rocket will blow out or not, the temperature range 
of this uncertainty being close to 20 F. Since in 
forward firing the sight setting is influenced very 
markedly by the burning time, this would mean 
that in this 20-degree range one would not know 
what sight setting to use and, if the wrong guess 
were made, the rocket would be too inaccurate to be 


useful. On the 5.0-in. and 11.75-in. aircraft rocket 
motors, therefore, 110 F was chosen as the tem- 
perature at which half the disks blow out, this tem- 
perature being above that normally required in 
practice and having a short burning time so that 
the error in gravity drop caused by a disk’s blowing 
unexpectedly would not be so great as at lower 
temperatures. Thus the blowout disk has been used 
primarily as a safety valve rather than as a tem- 
perature range extender. Its effect on the tempera- 
ture range is striking, nevertheless, particularly on 
the 5.0-in. HVAR, which operates with a very small 
percentage of failures at 140 degrees. Its lower 
temperature limit is not known, but it has been fil ed 
successfully after having been packed in solid carbon 
dioxide (sublimation point —109 F) over night. 

Blowout disks have been made of annealed copper 
because a copper disk is thicker for a given blowout 
pressure than a steel one, and hence small variations 
in thickness have less effect on the blowout pressure. 
A disk fails by first bulging out into a hemispherical 
shape and then shearing. Empirically it has been 
found that the failure pressure cold (i.e., in a testing 
machine) can be calculated fairly accurately if a 
shear strength of 25,000 to 26,000 psi is used. The 
mean blowout pressure measured in static firing is 
only 5 or 10 per cent higher than that. 

Since rockets with blowout disks are usually de- 
signed with relatively large nozzle K’s, they will be 
just as unsafe at high temperatures if the disk 
should fail to blow as if too small a nozzle had been 
used. The easiest ways to make an error in this 
regard appear to be the substitution of too thick a 
disk or the inclusion of two disks. Because of the 
importance of having the proper disk, it has 
seemed desirable to eliminate any possibility of 
error by two provisions: 

1 . Each disk is gauged for thickness and its thick- 
ness marked on it with a rubber stamp in a position 
where it is visible from the outside of an assembled 
motor and can be checked by the final inspectors 
and the loading crew. 

2. The “disks” are made cup-shaped rather than 
flat so that the ring or grid stool which holds them 
in place in the nozzle plater cannot be properly 
assembled if two disks have been used. If two 
rockets were made with blowout disks of different 
thickness, it would be desirable to make them 
completely noninterchangeable by a similar trick. 

Proper insulation of the disk from the motor 
gases is obviously essential to its proper function. 


TAILS 


261 


Asbestos and fiberboard have been used success- 
fully, but the best insulation is probably a molded 
disk of asbestos-filled bakelite. The insulation 
should fit snugly but not tightly in its hole and be 
completely covered with a fireproof hard-setting 
plastic material (like Permatex No. 2) to eliminate 
any gas leakage around the edges. If the insulation 
for a blowout disk of small diameter is a press fit 
in its hole, it may resist being ejected so that an 
erratic increase in blowout pressure results. 

234 TAILS 

23,41 Types of Tails 

Fins for rockets show an extreme diversity of size 
and type. Those on the Army 4.5-in. rocket, for 
example, are 4 in. long and % in. wide, whereas 
those on the CIT target rocket measure 18 by 36 in. 
The target rocket is, of course, a special case, since 
its fins were made as large as possible for visibility, 
whereas ordinarily we wish to make fins as small as 
is consistent with adequate stability. Target rocket 
fins are, therefore, treated as a separate problem in 
Chapter 18, and in this section it will be assumed 
that the primary function of a fin is to stabilize the 
rocket in flight. 

The design of a tail n is always a compromise 
between a number of mutually contradictory re- 
quirements. Thus accuracy requires large tails, 
whereas space and weight considerations and air 
drag favor small. Simplicity and cheapness of 
manufacture favor fins made from a single thickness 
of metal and welded to the tube, whereas weight 
and convenience may favor double fins with rela- 
tively complicated attaching mechanisms. The 
usual factors controlling fin design include: 

1. Adaptability to the type of launching con- 
templated. 

2. Accuracy. 

3. Strength. 

4. Air drag. 

5. Shipping space. 

6. Sometimes provision of electrical contact. 
Two principal types of tails have been used: ring 
tails and fin tails. 


n For brevity we shall use the term “tail” instead of “fin 
assembly” in referring to the aggregate of the fins, rings, or 
members of whatever shape which constitute the stabilizing 
member. 


23,4,2 Ring Tails 

Ring tails haVe been used on most rockets having 
heads of larger caliber than their motors because, 
except in the unusual case of the ASR where the 
center of mass of the round is in the head, it is neces- 
sary to have a ring behind the center of mass with 
the same diameter as the head in order to fit a 
simple launcher. To provide easy electrical contact, 
the tail has always consisted of two rings, one 
attached to the motor tube and the other insulated 
from it. This combination makes the ring tail 



c 


Figure 11. CWR tails: (A) plain ring tail, (B) 
ring tail with radial fins (CWR-N), (C) experi- 
mental ring-and-fin tail designed for maximum 
stability. 

rockets adaptable to very rapid loading in any 
orientation such as is required in combat and to 
automatic launching. The design is inherently 
strong, and the thickness of metal used is deter- 
mined by the rough treatment experienced in 
handling and in being jammed against the knife 
edges of the electrical contacts, rather than by the 
relatively small aerodynamic forces encountered in 
flight. For withstanding water entry and for 
stabilizing underwater trajectories, ring tails are 
very satisfactory. 

As used on CIT rockets, the ring tails are not 
very efficient in stabilizing the rounds because the 
ring diameters are relatively small and not much air 


262 


MOTOR DESIGN 


passes through them, especially at high velocities, 
because of the shielding by the head. Usually the 
accuracy of the rockets was adequate for the tactical 
situation, but in the case of the CWR, radial fins 
extending beyond the ring were finally added, as 
shown in Figures 11 A and B, to increase the 
stability and decrease dispersion. High-speed 
water tunnel tests indicated that this change would 
reduce the yaw oscillation distance a from 236 to 
192 ft, whereas field firings gave values of 260 and 
215 ft. The water tunnel tests showed also that a 
further decrease of a to 166 ft could be made by 
moving the tail back about 1 caliber as shown in 
Figure 11C and that considerably more water 
passed through the ring under these conditions. 
Whether the same would be true when the effects 
of the rocket jet are added is problematical, but it 
may be possible to increase the efficiency of ring 
tails by thus moving them back and still retain the 
advantage that no part of the tail projects beyond 
the head diameter. 

23 4 3 Fin Tails 

Fin tails have been used on all the aircraft rock- 
ets, and again the choice was dictated by the 
launching method. Since the motor and head were 
of the same diameter, 0 fin tails allowed the rockets 
to be attached closer to the airplane with a conse- 
quently smaller drag. The width of the fins (i.e., 
in the radial direction) was also determined by the 
space limitations, and in all cases the length was 
made approximately V /2 times greater than the 
width in order to obtain the requisite strength. 
This ratio of length to width appears to be a good 
one, at least for subsonic rockets. p 

For rockets small enough to be handled manually, 
the forces encountered in handling are again the 
determining factor in the strength. In sizes com- 
parable to Tim, however, it ceases to be practicable 
to make fins so strong that they will support the 
weight of the rocket, and the aerodynamic forces 
are determining. These forces are difficult to cal- 
culate, but are not large in practice because appre- 
ciable yaws are obtained only at low velocity. 

Economy in shipping space demands that the fins 
be detachable from the motor. In practice this 

0 The 5.0-in. AR with the 3.25-in. motor is an exception to 
this rule, but there the controlling factor was the use of a motor 
already in production. 

p See discussion under HVAR in Chapter 19. 


means that fins for large rockets which cannot be 
boxed in groups of four with the fins nested between 
them (as was done with the 3.25-in. AR motor) 
will have fins individually detachable, and hence, 
to accommodate the locking mechanism, the fins 
may be made of two pieces of metal, dished so as 
to leave a space between. The double fins also have 
the advantage of being relatively strong with thin 
metal. On the British RP-3, detachability was 
achieved with remarkable simplicity and effective- 
ness with a single-thickness fin. Although it is 
almost certainly the best rocket fin in existence, it 
could not be copied in CIT rockets because slots in 
the motor tubes were not permissible. 

Drag was mentioned as a factor in fin design, but 
nothing more has been said about it, and in fact 
little consideration was given to it in the design of 
CIT rockets. The reason is that for short-range 
firing, such as aircraft forward firing, there is little 
to be gained by small reductions in the drag, since 
the total drag is large but its effect is small. For 
long-range rockets, this would not be true. 

Another thing about fin design may be conspicu- 
ous by its absence — namely, any mention of folding 
fins. These have been used effectively on the Army 
4.5-in. rocket, but were not tried by CIT. The 
reason is simply that, since integral formed nozzles 
cannot be used with the thin- walled rockets, there 
is no place to put the Army type which opens back. 
Fins which open out sideways are subject to serious 
objections: (1) inaccuracy, since the first moment 
after launching is the time when stability is most 
needed, and (2) practical difficulties in making a 
foolproof latch to hold them in the open position. 

23 5 SUSPENSION LUGS 

Suspension lugs and lug bands have been used 
only on forward- and backward-firing aircraft rock- 
ets where the air drag of the launcher is a prime 
consideration. In almost any other conceivable 
application, the drag of the rocket itself would be of 
greater importance, and lugs would be omitted. 
The drag of the front lug, in particular, can be quite 
significant because it is placed in a portion of the 
rocket which would otherwise usually be aero- 
dynamically 1 ‘clean,” and its presence thus increases 
the turbulence along almost the full length of the 
rocket. 28 

The shape of the lugs being dictated by the shape 
of the rocket and the method of launching, little 


MOTOR SEALS AND GRAIN SUPPORTS 


263 


can be said about it in general. It should be noted, 
however, that the shape now standard on the 3.25- 
in. and 5.0-in. aircraft rocket motors is certainly 
not ideal, resulting as it did from the historical 
accident that long T-slot launchers were already in 
combat use before the advantages of post launchers 
(or “zero-length” launchers) were established. The 
front lug is not very strong, is difficult to manu- 
facture, and has more drag than would be desired. 
The ideal would probably be two lugs side by side 
about 90 degrees apart on the front and one be- 
tween them on the rear, so that the rear lug could 
be made higher and stronger than is now the case 
and still not interfere with the front post when it 
passes. This type of suspension was used for ex- 
perimental aircraft firings of Tiny Tim from fixed 
wing launchers. 

Whenever their use is possible, welded lugs are 
much preferable to lug bands because (1) they 
assure that the position and spacing is always cor- 
rect, (2) they are easier to make strong enough, and 
(3) they are cheaper to manufacture. 


23 5 1 Strength of Lugs 

The basic data for determining the required 
strength of a lug is usually in the form of the maxi- 
mum values of yaw, roll, and pitch which the air- 
craft is expected to undergo in the most extreme 
maneuvers contemplated or possible. The transla- 
tion of these specifications into forces in various 
directions on the lugs is obvious and straightforward 
and typically results in strength specifications which 
are difficult to meet. Fortunately, the basic data 
by their very nature contain a considerable safety 
factor, so no additional factor need be interjected. 
For carrier-based planes, there is also a specification 
of the maximum fore-and-aft accelerations en- 
countered by the airplane in catapulting and 
arrested landings, but the maximum fore-and-aft 
force which the rocket itself will experience usually 
depends upon vibration, in the case of wing- 
mounted rockets, and its magnitude is difficult to 
estimate . 

Most of the difficulties with lug bands are elim- 
inated or greatly reduced if the bands can be made 
tight enough. In the case of the 11.75-in. motor, 
“tight enough” meant going to specially heat- 
treated high-tensile steel. The best design for the 
tightening mechanism on a lug band is probably 


that shown in Figure 12, which was adopted for the 
5.0-in. and 11.75-in. motors after experience with 
several other types. In case slippage along the tube 
is undesirable, as it is for the post launcher where 
only one post contains a latch, it can be eliminated 
by drilling a shallow flat-bottom hole in the motor 
tube and having a pin on the lug band which pro- 
jects into it. This was done on the nonwelded CIT 
design of the HVAR (5MA4). (See Section 19.4.2.) 
The drill marks on the 3.25-in. Mk 7 motor tube 
served to position the lug bands when they were 
attached but did not significantly reduce the slip- 
page because of their tapered sides. 




Figure 12. Final design of lug band clamp. 

23 6 MOTOR SEALS AND 

GRAIN SUPPORTS 

Two factors in motor design have not yet been 
mentioned. That of making electrical contact to 
the igniter is a rather simple and specific problem. 
It is mentioned in Chapters 18 and 19 in connection 
with the 4.5-in. BR, the 3.25-in. AR, and the 11.75- 
in. AR motors, but it will not be discussed here in 
detail. 

The problem of supporting the grain at the front 
end and sealing the two ends of the rocket against 
moisture are, on the other hand, practically iden- 
tical for all rockets, except for scale effects. 


23,6 1 Grain Support 

In order to eliminate the possibility of cracking 
the propellant grain as a result of impact against 
the grid when the igniter fires, it has been con- 
sidered desirable to hold the grain at the extreme 
rear end of the motor by means of some type of 


264 


MOTOR DESIGN 


grain support at the front end. For small-diameter 
motors, where the weight of the grain is small and 
its length short, the front motor seal is adequate for 
this purpose. It is simply pushed in until it seats 
firmly against the igniter which in turn contacts the 
grain. When the length of the grain exceeds about 
2 ft or its weight becomes of the order of 10 lb, this 
simple procedure is not adequate. If the grains are 
not thoroughly annealed, they shrink with age as 
the strains introduced during extrusion are relieved. 
Temperature changes also cause changes in length 
which can be significant on very long grains, and it 
is desirable to have something to take up these 
length changes without allowing the grain either to 
become loose or to exert so much force on the front 
sealing disk that the seal is broken. The best sub- 
stance which has been found for doing this is a thick 
felt disk compressed to about % or % of its uncon- 
fined length. Felt has no undesirable effects on the 
propellant, nor is it affected by the propellant 
fumes. Felt disks are used in both the 3.25-in. and 
5.0-in. aircraft rocket motors. 

With the heavier grains, the accelerations ex- 
perienced during handling might be large enough 
that the grain would move the front seal if it were 
not reinforced. In the 3.25-in. and 5.0-in. motors, 
this support is provided by the front thread pro- 
tector. The 11.75-in. motor is special in that the 
grains are held against the grid by the charge sup- 
port independently of the motor tube. 

It is interesting that the only test which was 
made of the necessity for holding grains firmly 
against their grids showed that it was not necessary. 
Two rounds of the 5.0-in. HVAR were fired at 120 F 
with 20-lb heads which would give them an accelera- 
tion of more than 80$. The rounds flew normally, 
although the grains, with grids attached, were 
separated from the grid stools by distances of 3M 
and 4 in. 17 Despite this evidence, the require- 
ment that grains be firmly seated is based on good 
logic, particularly since in some cases the grid can 
rotate if it becomes loose, a circumstance which 
would almost certainly cause a motor burst. 

23 * 62 Seals 

The first seals used by CIT to keep moisture out 
of motors were binderboard and fiberboard disks 
pressed into position and sealed with glyptal lacquer. 
When tests had demonstrated that such disks did 
not in fact keep out moisture if the motors were 


subjected to extreme temperature changes, cellulose 
acetate was substituted, and it in turn was found to 
be inadequate and displaced by steel. The complete 
story of the tests made to determine the best seals 
is contained in reference 29, and will therefore be 
merely summarized here. 

No nonmetallic materials were found which, 
either in the shape of disks or cups, would protect 
motors from extreme conditions of exposure . Fiber- 
board absorbs water through even the best paint 
seals, swelling up and softening. Thermoplastics 
have a thermal coefficient of expansion much 
greater than that of steel, and as a result plastic 
closures shrink away from the motor tubes when 



1 INCH 


Figure 13. Front end motor seals for 5.0- and 
3.5-in. motors. 


MOTOR SEALS AND GRAIN SUPPORTS 


265 





Figure 14. Metal nozzle seals: (A) SCAR; (B) HVAR pigtail seal (for one nozzle) ; (C) HVAR plain seal 
(for other 7 nozzles) ; (D) BR, CWR, and similar motors with electrical contacts on tail use seal similar to 
(C) but with two wires brought around seal edges at opposite sides as shown; (E) 3.5-in. spinner. 


266 


MOTOR DESIGN 


they are cooled, breaking the paint seal at the 
edges . Bakelite is too brittle to be inserted tightly 
without cracking. 

Of all metal closures, steel seems to be unquestion- 
ably the best. Brass closures corrode rapidly in 
contact with steel motor tubes because of electro- 
chemical action. Aluminum corrodes even more 
rapidly, large holes being eaten away leaving the 
covering film of glyptal unsupported, and in addi- 
tion it is too soft so that it was difficult to insert 
aluminum closures without deforming them. Steel 
closures are easy to construct, can be inserted 
rapidly, and, if they have the proper thickness so 
that they still have some spring in them after being 
inserted, they give effective protection against 
moisture even without a perfect paint seal. They 
have the same coefficient of expansion as the motor 
tubing and the nozzles. Some objections have been 
raised to them because of their missile hazard, but 
tests indicate that it is no more serious than with 
fiberboard disks. 

Three types of seals are required for a rocket 
motor. For the front motor seal, the most effective 
design appears to be a flat-bottomed cup either with 
plain or re-entrant sides as shown in Figure 13. 
The “blowout patch” in the center was evolved 
after considerable experimentation as the best de- 
vice for opening quickly at low pressures but still 
being easily moisture-proofed when it is in place. 
In some motors it serves the purpose of admitting 
the gas to the pressure-arming base fuze, and in all 
motors it assures that the motor would not become 
propulsive in case of accidental ignition when the 
head was not screwed on. 

For nozzle seals, a simple shallow cup with ta- 


pered sides is adequate when no wires must come 
through it. To accommodate the igniter leads, the 
cup must be made slightly more complicated as 
shown in the examples in Figure 14. Even when 
good nozzle seals are used, care must be taken that 
moisture does not enter through the nozzle threads 
(if any) or along the cotton insulation or filler in 
the igniter leads. 

For the 3.5-in. and 5.0-in. spinner motors, it 
appeared simpler to seal the nozzle end with a single 
metal disk instead of closing each nozzle separately. 
Thus the wires connecting to the contact rings are 
completely enclosed and protected. The standard 
nozzle end seal for the 5.0-in. motors is shown 
in Figure 4 of Chapter 20. The seal for 3.5-in. 
motors, shown in Figure 14E of Chapter 23, is 
basically similar but has a flange extending beyond 
the diameter of the round to keep it from sliding 
forward in tubular launchers. 

For sealing all these steel cups, the best material 
found is General Electric glyptal red lacquer No. 
1201, with the addition of 7 per cent by weight of 
aluminum powder, which toughens it and makes it 
dry better around wires insulated with nylon. 
Nozzle closures hold better if the edges are painted 
with thinned glyptal containing emery, 200-mesh 
being the optimum granulation. 

The larger motors are so expensive that extra 
precautions have seemed desirable to keep them 
dry, and auxiliary seals have been used at both 
ends. At the front, the extra seal is easily incorpo- 
rated in the thread protector, but the design of the 
rear one depends on the motor. Blowout patches 
may be required in these also to keep the motor 
from being propulsive when shipped. 


Chapter 24 

EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 

By C. W . Snyder 


241 INTRODUCTION 

I N THE FOLLOWING TWO CHAPTERS, We shall disCUSS 
briefly and qualitatively the exterior ballistics of 
rockets. It is an exceedingly large field and one of 
considerable complexity; we shall attempt merely 
to lay a groundwork for understanding why rockets 
behave in flight as they do and what methods are 
used to predict their performance, and to indicate 
where more thorough discussions of various aspects 
of the problem can be found. The theoretical basis 
of the subject is treated in detail in a book to which 
we shall refer frequently by the abridged title of 
Exterior Ballistics; 1 this is the source of much of the 
following material. 

Specification of a Rocket’s 
Motion 

It will be well to have clearly in mind at the out- 
set the precise meanings of the terms which will be 
used in the description of the sometimes complex 
motions of a rocket in flight and the symbols by 
which they will be denoted. 51 There is an important 
theorem of mechanics which states that the motion 
of the center of mass of a solid body which is acted 
upon by any arbitrary combination of forces is the 
same as if all the body’s mass were concentrated at 
that point and all the forces acted on that point. 
Consequently, the simplest way to treat the motion 
of a solid body and the way that is always adopted 
in practice is to consider first the motion of the 
center of mass and then independently of this motion 
to consider the rotations of the body about the 
center of mass. 

The path of the center of mass through space is 
called the trajectory of the rocket. It is in general a 
complicated curve, but the simplest case, when it 
lies in a vertical plane, is illustrated in Figure 1. 
When the rocket (represented by a small arrow in 
Figure 1) is at the point C , its center of mass is 
moving in the direction of the tangent to the tra- 

a The notation here is taken from Exterior Ballistics 1 and 
agrees in the main with that of earlier CIT reports. 


jectory (shown as a dashed line intersecting the 
horizontal coordinate axis). The angle 6 between 
the initial orientation of the rocket (i.e. , the launcher 
orientation) and the tangent to the trajectory at a 
particular time we shall for brevity call the trajec- 
tory deviation. In addition to 0, we must of course 
know the orientation of the plane of the trajectory 
(i.e., the plane containing the launcher line and the 
tangent to the trajectory) in order completely to 
specify the rocket’s direction of motion. The angle 
0 is the more important quantity, however, because, 



Figure 1. Trajectory of rocket in vertical plane. 

except for gravity, nearly all the forces acting on a 
rocket are unchanged when the orientation of the 
plane is changed. 

In general, the rocket will not be pointed in 
exactly the direction in which its center of mass is 
moving, and in this case it is said to have a yaw. 
The rocket can yaw m any direction, b but for tinners 
the usual case is that shown in Figure 1, where the 
yaw is in the plane of the trajectory. The yaw angle 
8 is the angle between the trajectory and the axis 
of the rocket. 

A third angle which is usually of less importance 
than either 0 or 8 is the rocket orientation angle 0. 
It is the angle between the rocket axis at any time 
and the line through the launcher. In the plane case 
shown in Figure 1 we have obviously the relation: 

<j> = 8 + 0 , 

b This usage of the term yaw is slightly different from the 
nautical usage. Thus a ship yaws sideways but pitches up 
and down. 


267 


268 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


but this will not hold in general unless we consider ated laterally . The reaction on the motor tube tends 
the angles as vectors, a complication which we will to damp the rotation. This so-called “jet damping 
avoid here. torque” is too small to be important in practice. 


242 FORCE SYSTEM OF A FINNER 


Because of its complex shape, both interior and 
exterior, a rocket is subject during flight to a multi- 
plicity of complicated forces, and an understanding 
of rocket motion requires that we replace this force 
system with a simpler one which produces the same 
accelerations and velocities. An elementary theo- 
rem of mechanics assures us that it is always pos- 
sible to do so. The resulting force system is, of 
course, arbitrary, and it is chosen to make the ana- 
lytic representation as simple as possible. In par- 
ticular, it is convenient to consider separately the 
forces arising from the combustion of the propellant 
and those arising from the presence of the atmos- 
phere, since the former disappear after the end of 
burning. 

24.2.1 The j et F orce all R Torque 

From consideration of the conservation of linear 
momentum, we derived in Chapter 21 the fact that 
the ejection of the propellant gas from the nozzle 
results in a force on the rocket which was called the 
thrust. For simplicity we shall assume that its 
direction and magnitude are constant throughout 
burning and that it ceases abruptly. Actually, of 
course, its time variation is given approximately 
by the pressure-time curve (see Chapter 21), but 
the assumption of constancy introduces fairly small 
errors, which are discussed in Exterior Ballistics. 1 

In the ideal case, the line of action of the resultant 
jet force would lie along the rocket’s long axis and 
pass through its center of mass. Since these condi- 
tions are never perfectly fulfilled, we obtain, in addi- 
tion to the forward thrust, a torque of magnitude 
equal to the product of the thrust by the distance 
between its line of action and the center of mass. 
This is the so-called “jet malalignment torque.” 0 

One other subtle torque results from the action 
of the gas on the rocket. If the rocket is rotating 
about a transverse axis during burning, the gas as it 
flows down the motor tube will have to be acceler- 

c Actually there may be two types of malalignment, 13 but 
only one is important in practice. 


Aerodynamic Forces 

The effect of air on the rocket in flight can be 
treated with sufficient accuracy by means of two 
forces and two moments, defined as follows. Con- 


v 



DECREASING YAW 

Figure 2. Aerodynamic forces and torques act- 
ing on fin-stabilized rocket. 

sider a projectile moving through still air in the 
direction of the vector V of Figure 2, having a yaw 
represented by the angle 5 and having a certain 
instantaneous angular velocity about the transverse 
axis perpendicular to the plane of the paper through 
its center of mass. Although the aerodynamic 
forces are produced by the distribution of pressure 
over the entire surface of the projectile, we need not 
consider the distribution in detail because its effect 
is the same as that of a suitable single force Fa act- 
ing at an arbitrarily chosen point (for convenience 
taken to be the center of mass) plus a suitably 


USE OF THE FORCE SYSTEM 


269 


chosen torque. If F A is resolved into components 
along and perpendicular to the trajectory (i.e., to 
the velocity vector V), the former is the “drag” 
F d and the latter is the “cross-wind force” or “lift” 
Fc . Of the total torque, the major part, which 
depends upon the yaw but not on the transverse 
angular velocity, is called the “righting moment” 
or “restoring moment” M since it tends to reduce 
the yaw; and the small part which varies with the 
transverse angular velocity is called the “damping 
moment” M D because it tends to reduce the angular 
velocity and momentum. In the figure it is assumed 
that the yaw is decreasing so that M D , tending to 
oppose the decrease, acts in the direction opposite 
to M . When the yaw is increasing, both moments 
tend to oppose the increase. The principal effect 
of the drag is to decrease the velocity and range of 
the rocket, whereas that of the righting moment is 
to stabilize the rocket and to produce oscillations 
in the orientation of the rocket whenever it yaws. 
The cross force and the damping moment are of 
relatively minor importance and serve chiefly to 
damp the oscillations. It was noted in Chapter 21 
that a righting moment exists for small yaws only 
if the fins are sufficiently large, and that F D , Fc, 
and M are nearly proportional to the square of the 
velocity V, up to about 800 fps. Hence we set 


F d = mV 2 c; 

a) 

M — nV 2 sin 8 « fxV 2 8; 

(2) 


where c is the deceleration coefficient, m the mass, 
and /a the righting moment coefficient. Equations 
(1) and (2) are equivalent respectively to equations 
(16) and (22) in Chapter 21. 

The force system of Figure 2 is not, of course, the 
only one that will produce the same acceleration 
of the rocket as does the actual pressure distribu- 
tion. It is possible to find a point on the axis of 
the rocket such that the resultant force F A applied 
at this point gives the moment M and hence is fully 
equivalent to the entire aerodynamic pressure dis- 
tribution. This point is called the center of pressure, 
and the force system is shown in Figure 3. It is 
more convenient than that of Figure 2 for visual- 
izing the effect of aerodynamic forces but less useful 
for computation. The center of pressure must lie 
to the rear of the center of mass in order for a finner 
to be stable. 

If the rocket is traveling through water or earth, 


the aerodynamic forces are replaced by a different 
force system, which will be discussed later. 

24 2 3 Other Forces 

To complete the list of forces which determine a 
rocket’s trajectory, the pull of gravity and the 
reaction of the launcher must be included. The 
latter is effective for such a short time that it can 
be considered as an impulse. 


DIRECTION OF MOTION 
WITH RESPECT TO AIR 



Figure 3. Alternative aerodynamic force system. 


243 USE OF THE FORCE SYSTEM 

Through an analytical representation of these 
various forces and torques, it is possible to set up 
the equations of motion of a rocket in flight, both 
during and after burning, and, at least theoretically, 
to solve them for the motion of the rocket under 
various initial conditions. This analysis is devel- 
oped in detail in Exterior Ballistics. 1 In practice, 
of course, the solution of the equations is extremely 
difficult unless a number of simplifying assumptions 
are made. 


270 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


244 RANGE OF A 

GROUND-FIRED ROCKET 

The quantity of first concern is usually the range 
of the rocket or, more accurately, the mean range 
of a large number of identical rockets fired under 
the same conditions. For this calculation, one 
assumes that the thrust, the drag, and the pull of 
gravity are the only forces acting. We shall see that 
the solution of even this much simplified case is 
very difficult unless the velocity is small. 

The vacuum range X of a projectile in free flight 
after launching at velocity Vo and elevation angle 
do was given in Chapter 21 as 

x _ Vo 2 sin 2fl 0 . ^ 

It was noted that this requires modification for 
rockets on two counts: (1) it must be corrected for 
the burning time, since the rocket is not in free 
flight until after the jet force ceases, and (2) the 
effect of aerodynamic forces cannot in general be 
neglected. Even as slow and dense a rocket as 
the “Mousetrap” antisubmarine rocket [ASR] (175 
fps) attains only 95 per cent of its maximum vac- 
uum range. 

The effect of burning time is most conveniently 
allowed for by the concept of the “equivalent 
shell.” d An equivalence is set up between the 
rocket and a hypothetical shell which have coinci- 
dent trajectories after the rocket stops burning. 
Thus, having translated the initial conditions of 
projection of the rocket into those of the equivalent 
shell, we can use equation (3) or other more exact 
relations from shell theory to determine the range 
and trajectory of the rocket subsequent to burning. 
The expressions for accomplishing this translation 
are as follows. 

If a rocket is fired at a quadrant elevation greater 
than zero, its velocity at the end of burning will fall 
short of that calculated from momentum considera- 
tions [equation (6) of Chapter 21] for two reasons: 
there is a component of gravity acting backwards 
along the trajectory and the air resistance is con- 
tinuously removing energy during the acceleration. 
The actual “burnt velocity” will be, instead of V Q 
to be expected in a vacuum, 

V b = Vo — t h (g sin do + \cV b 2 ) , (4) 

d The theory of the equivalent shell is worked out in ref- 
erences 2 and 3. 


in which t b is the burning time (duration of thrust). 
The factor }/& takes care of the fact that we should 
actually use some kind of average velocity during 
the burning period instead of V b itself in computing 
the effect of air resistance. These same two effects 
will reduce the velocity of the equivalent shell, but 
they will have only half as long a time to act, since 
the rocket, starting from zero velocity, has during 
the burning time an average velocity half that of 
the shell. Hence the shell, if it is to match the 
rocket flight in space and time, will be fired later 
than the rocket by half the burning time and will 
have an initial velocity 

V e = Vo - it b {g sin do + \cV b 2 ) 

= V h + \t b (g sin 0 O + \cV * 2 ). 

Finally, because of the greater gravity drop of the 
rocket, the equivalent shell will be fired at a lower 
angle of elevation than the rocket by an amount 
proportional to the length of time which the rocket 
burns beyond the launcher. In fact, the elevation 
angle of the equivalent shell will be 

d e = do — ^y-(tb — t p ) cos do (in radians), (6) 

where t p is the time spent on the launcher (desig- 
nated by the subscript p because at the time this 
theory was developed, the term “projector” was 
in vogue) . 

24.4.1 Air J) ra g 

It is immediately evident that no ballistic calcu- 
lations can be made without a knowledge of the 
value of the drag of the rocket at all velocities that 
it attains. Aerodynamics has not progressed to 
the point where the drag coefficient of a projectile 
can be computed on purely theoretical grounds, 
but by a combination of theory and empirical results 
it is possible to make surprisingly close estimates 
of the drag coefficient of an aerodynamically clean 
projectile. However, if it has large lugs, fin braces, 
or other irregular projecting parts that tend to pro- 
duce large contributions to the drag, the estima- 
tion is much more difficult. Examples of such 
calculations are given in Exterior Ballistics 1 and in 
references 4 and 5. 

The method employed is to divide the total drag 
into five parts: 

1. Head resistance. 

2. Base drag due to reduced pressure at the rear. 


RANGE OF A GROUND-FIRED ROCKET 


271 


3. Skin friction of the cylindrical motor tube. 

4. Skin friction of the fins. 

5. Drag due to lugs and other irregularities. 
Parts 1 and 2 can be estimated from the known drag 
of a shell having a nose shape as much as possible 
like that of the rocket under consideration. The 
skin frictions, 3 and 4, are calculable from aerody- 
namic theory. When the contribution of 5 is not 
significant, the sum of parts 1 to 4 is often in fairly 
good agreement with the experimentally determined 
drag. An interesting and important example in 
which this is not the case is analyzed in Table 1. 


Table 1. Relative contributions to total drag for 3.5-in. 
aircraft rocket (CIT Model 5) . 


Velocity 

(fps) 

Head 
resistance 
and base 
drag 

Motor 

skin 

friction 

Fin 

skin 

friction 

Unac- 

counted 

for 

C{V) 
C (600) 
(experi- 
mental) 

600 

32% 

18% 

34% 

16% 

1.00 

800 

33% 

17% 

33% 

17% 

1.00 

1,000 

39% 

15% 

30% 

16% 

1.06 

1,200 

37% 

8% 

15% 

40% 

2.05 

1,400 

46*% 

74% 

144% 

34% 

2.07 


Columns 2, 3, and 4 give the theoretical estimates 
for various parts of the drag, expressed as percent- 
ages of the total experimentally determined drag. 
Column 5 gives the percentage of the total drag 
which the theoretical analysis does not account for. 
Presumably most of this discrepancy is caused by 
the unusually large lug bands which the motor car- 
ries in order to accommodate 5.0-in. as well as 
3.5-in. heads. Column 6 gives relative values of 
the total deceleration coefficient at various veloci- 
ties. Theoretically the increase in drag between 
low and high velocities should be approximately 
3 to 2 rather than 2 to 1 as actually observed, illus- 
trating the well-known fact that good streamlining 
is much more important above sonic velocity. 

24.4.2 Calculation of Range 

If one has precise knowledge of the value of the 
deceleration coefficient as a function of velocity, 
it is theoretically possible to make accurate trajec- 
tory and range calculations by means of numerical 
integration, but the labor involved makes such 
calculations impracticable except on a modern me- 
chanical or electrical integrator, few of which are 
in existence at present. Complete ballistic tables 
have been worked out for several different shell 


shapes, and what is done in practice is to pick the 
one of these standard drag functions which most 
nearly approaches that of the rocket in question 
and use the ballistic tables corresponding to that 
function. 

This method of calculation is quite satisfactory 
for low-velocity rockets, i.e., in the velocity range 
where the deceleration coefficient can be assumed 
constant. For firings at quadrant angles below 15 
degrees or for segments of a trajectory in which the 
direction of the trajectory does not change by more 
than about 30 degrees, the Didion-Bernoulli meth- 
od lbt 6 is probably the most convenient. For rockets 
fired from the ground at higher quadrant elevations, 
the Otto-Lardillon tables 7 have been reduced to 
more convenient graphical form in reference 2 and 
have been found to be sufficiently accurate and 
very useful for predicting ranges. 

The greatly increased complexity of the problem 
at higher velocities arises from the varied shapes of 
rockets. Skin friction, the turbulent drag of the 
projections, and the other factors will contribute 
to the total drag in varying proportions for various 
rockets, and each factor will, in addition, vary with 
velocity in a different manner. Thus no one drag 
function can be expected to be a sufficiently good 
approximation for more than a very restricted fam- 
ily of rockets. Several resistance functions have 
been found useful in particular rocket ballistics 
problems, the one most frequently used being that 
of the French Commission de Gavre, not so much 
because of its merit but because most of the avail- 
able ballistic tables (in particular those usable for 
high-angle fire from the ground) are based on it. 
The Gavre function is based on drag measurements 
of an obsolete type of shell having a relatively blunt 
nose and no boattailing, and the fact that many 
contemporary rockets have these same character- 
istics provides some justification for its use. 

It would take us too far afield to discuss the 
details of the methods of range calculations. They 
are given in Exterior Ballistics . lb In addition, a 
good bibliography on the subject is contained in 
Rocket Fundamentals . 8a 

2443 Launcher “Tip-off” Effects 

In correcting the quadrant elevation of the rocket 
to that of the equivalent shell [equation (6)], it 
was assumed that the “tip-off” is negligible. During 
the short time when the center of gravity of the 


272 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


rocket is off the launcher but the tail (or rearmost 
point in contact with the launcher) is still in con- 
tact, a torque exists tending to give the rocket an 
angular momentum about a horizontal axis. In the 
case of spin-stabilized rockets, the combination of 
this torque with the gyroscopic effect results in a 
deflection of the round to the left (assuming right- 
hand spin), but for a fin-stabilized rocket, tip-off 
simply reduces the effective launching angle and 
hence reduces the range if the quadrant elevation 
is 45 degrees or less. Theoretical analysis e shows 
that the amount of rotation during tip-off is negli- 
gible, but that the angular velocity imparted to the 
rocket persists and continues to decrease the effec- 
tive elevation angle so that, for rockets which are 
launched at very low velocities, the total reduction 
in effective launching angle can amount to several 
degrees. 

Tip-off can be reduced in two ways: 

1. By reducing the ratio of burning distance to 
launcher length, either by using a longer launcher 
or a grain giving shorter burning time, so that the 
rocket is launched at higher velocity; or 

2. By arranging that the rocket is not constrained 
after the center of gravity leaves the launcher. This 
has been accomplished in certain cases by using a 
special launcher such as the “zero-length” launcher 
or, for the antisubmarine rocket, by making the 
diameter of the tail smaller than that of the head 
so that the tail does not touch the launcher at all. 

245 WIND EFFECT 

The effect of a uniform wind on the flight of a 
finner follows simply from the fact that the aero- 
dynamic moment is a righting moment. From 
whatever direction the wind is blowing, its force, 
being greater on the tail than on the nose, will push 
the tail downwind so that the rocket will head into 
the wind. This effect is most striking in the case of 
rockets fired from aircraft, i.e., in high relative 
winds. To take an extreme example, suppose that 
a 5.0-in. high-velocity aircraft rocket [HVAR] is 
fired from an airplane traveling 450 mph pointing 

e The theory of tip-off is derived in two local CIT reports 9 ’ 10 
for rockets like the 4.5-in. barrage rocket in which a single 
point on the tail touches the launcher after the head leaves it, 
and in Rocket Fundamentals , 8 for rockets of uniform diameter 
for which the point of contact with the end of the launcher 
moves along the rocket. Both cases are treated in Exterior 
Ballistics. 1 


10 degrees away from the resultant wind. Then, if 
the temperature is low so that the burning time is 
1.2 seconds or more, its apparent launching direc- 
tion at the end of burning will deviate from the 
wind direction not by 10 degrees but by less than 

0.2 degree. In ground firing, the effect is qualita- 
tively similar but much smaller. Consider a wind 
blowing at right angles to the launcher; then it is 
obvious that its effect is divided into two parts: 

1. During the burning period the action of the 
wind on the fins will turn the nose into the wind, 
and the jet will push the rocket in the direction that 
it points. As long as burning continues, the tangent 
to the trajectory, although oscillating slightly as 
shown in Figure 4, deviates on the average farther 
and farther from its original direction. If a long- 
burning rocket is launched at very low velocity, it 
may be pointing almost directly upwind, when it 
ceases burning, but, in the usual ground-firing case, 
the turning into the wind amounts to only a few 
degrees. 

2. After burning, the rocket will drift downwind. 
This drift comes about not, as might be thought, 
because of the cross-wind force (“lift”) but because 
of the action of the downwind component of the 
drag. The reason is that the period of oscillation 
of the rocket is small compared to the total time of 
flight, and, since the yaw oscillations are about the 
position in which the yaw and lift are zero, the 
effect of the lift approximately averages to zero . If 
the rocket is headed almost directly upwind, then 
obviously the only effect of the wind is to slow it 
down. In the usual case where the trajectory makes 
a large angle with the wind, the tangent to the tra- 
jectory turns gradually back toward its original 
direction and may go beyond it if the flight con- 
tinues long enough. Whether the resultant deflec- 
tion is upwind or downwind depends upon the ratio 
of total flight time to burning time, that is, upon 
the quadrant elevation and the temperature. 

The theoretical expressions for the deflection of 
the trajectory by a cross wind during burning are 
derived in reference 11 and in Exterior Ballistics f 
and the results are shown graphically in Figure 4 
in terms of dimensionless parameters which can be 
applied to any fin-stabilized rocket. For any given 
rocket, the ordinates are proportional to the deflec- 
tion of the trajectory from the original launcher 
line, the abscissas are proportional to velocity, and 
the parameters characterizing the various curves 
are proportional to the square root of the launcher 


WIND EFFECT 


273 



length. The symbols and their relations are as fol- 
lows: 

f = “velocity parameter.” 



= “characteristic function for trajectory deviation by a 
cross wind.” The actual angle 0 in radians of devia- 
tion of the tangent to the trajectory from its original 
direction for any particular rocket is obtained by 
substituting the proper values into the relation: 

e = 2?er. 

* (T 


a = distance traveled by the rocket during one cycle of 
yaw oscillation (ft). 

G = acceleration of the rocket (ft/sec 2 ). 
p = length of the launcher (ft) . 

V = instantaneous velocity of the rocket (fps). 

V p = velocity with which the rocket leaves the launcher 
(fps). 

V a = velocity of the rocket at the end of the first yaw oscil- 
lation cycle (fps). 

v <r = V2 g7. 

Wn = component of wind velocity perpendicular to the 
launcher (fps). 


274 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


By inserting into the graph the proper value of 
we can calculate the trajectory deviation at any 
time during burning or at the end of burning. 

To see the importance of the wind effect in ground 
firing, let us consider the effect of wind on the 
4.5-in. barrage rocket at 70 F, and at the end of 
the burning period. For this 

£ = 960 ft/sec 2 ; 

<7 = 265 ft; 

7(7 = 715 fps; 

p = 5 ft; 

V2fp« 0.2; 

Vb= velocity at end of burning = 355 fps; 
at end of burnings 0.5. 

Reading the value @w = 0.5 from the graph (the 
negative sign simply means that the deflection is 
upwind), we calculate for the trajectory deflection 
per unit cross wind the value of 0.7 mils per fps. 
Since the lateral dispersion (mean deviation) of the 
rocket is about 45 mils, it will apparently take a 
rather large side wind to change the center of impact 
by an amount comparable to the dispersion, espe- 
cially since part of this deflection is canceled out by 
the drift after burning. The actual effect of wind 
on the impact point of the barrage rocket is given 
in Table 2. f 


Table 2. Wind deflection of impact point for 4.5-in. 
barrage rocket. 


Propellant 

temperature 

m 

Angle of 
elevation 
(degrees) 

Lateral deflection 
into the wind for 
wind of 1 mph 

(yd) 

Increase in 
range for 
tail wind 
of 1 mph 
(yd) 

40 

20 

0.7 

1.6 

40 

45 

1.4 

1.6 

70 

20 

0.2 

1.2 

70 

45 

0.5 

1.7 

100 

20 

-0.1 

1.0 

100 

45 

-0.1 

1.6 


The effect on the trajectory of the component of 
wind in the direction of the launcher is virtually 
negligible, so that the general case of wind in any 
direction is obtainable from the curves of Figure 4. 
Thus the lateral deflection on the horizontal plane 
is obtained by inserting the component of wind 
perpendicular to the line of fire and dividing the 
result by the cosine of the quadrant angle. The 
f A more complete table is included in reference 12. 


change in effective launching angle by up-range and 
down-range winds is given by using as Wn the 
along-range component multiplied by the cosine 
of the quadrant angle. 

The calculation of the drift after burning can be 
done by simple methods familiar in artillery theory. 
They are discussed in references 11 and 13. 

246 TRAJECTORIES OF ROCKETS 
FIRED FORWARD FROM AIRPLANES « 

By far the most extensive application of external 
ballistic theory to rockets has been in connection 
with forward firing from aircraft, for it is only in 
this use that fin-stabilized rockets are sufficiently 
accurate to warrant accurate calculations of trajec- 
tories. As in the case of ground trajectories, the 
solution requires setting up the equations of motion 
of the rocket in the air and integrating them, but 
the solution is simpler here because of the relatively 
short flight times that have been used in practice. 
The methods of calculating the trajectories and the 
sighting tables required for various aircraft and 
various firing conditions are worked out in detail 
in reference 15 and in Exterior Ballistics. 1 

The characteristics of the rocket trajectory can 
best be understood through a comparison with those 
of the more familiar machine gun bullet . h If firing 
conditions are identical, the two trajectories differ 
mainly in the following three respects: 

1. Rockets are slower. The velocity of the fastest 
rocket used at present in forward firing is approxi- 
mately 1,350 fps, which is about half that of a 
.50-caliber machine gun bullet. Furthermore, it 
takes the rocket a relatively long time — of the order 
of 1 second, more or less depending upon the tem- 
perature — to reach its maximum velocity, whereas 
the bullet has its maximum velocity as it leaves the 
muzzle. The consequent longer time of flight of the 
rocket to a given range means that allowances for 
target speed and wind are much greater than in the 
case of machine gun fire. 

2. Rockets tend to follow the direction of flight 
of the aircraft , whereas bullets travel in the direction 
of the gun. The bullet travels close to the direction 
of aim because its muzzle velocity is so much greater 

g A basic reference on this subject is Firing of Rockets from 
Aircraft , 14 one of the CIT final reports under OEMsr-418. 

h See reference 16 for an excellent simple discussion of the 
genera] discussion of the general features of aircraft rocket 
trajectories. 


TRAJECTORIES OF ROCKETS FIRED FORWARD FROM AIRPLANES 


275 


than the speed of the airplane that the effect of 
the latter upon the motion of the projectile is rela- 
tively slight. We have already seen that the fins 
of a rocket, on the other hand, tend to align it with 
the airflow resulting from the combination of the 
velocities of rocket and aircraft. Since the launch- 
ing speed is low, the rocket is quickly aligned almost 
in the direction of the line of flight of the aircraft. 
This deflection toward the flight path is greater 
the less the launching speed of the rocket, and is 
almost 100 per cent with post launchers. 

3. The rocket trajectory is characterized by con- 
siderable curvature compared with the flat trajectory 
of a bullet. Not only does the longer time of flight 
lead to a greater gravity drop, but, in addition, as 
the rocket falls, the fins tend to align it along the 
trajectory so that there is also a component of the 
jet force downward contributing to the normal 
gravity drop. The consequent large curvature 
means that the sighting allowance required in aim- 
ing and its variation with dive angle are much 
greater for rockets than for guns. 


Trajectories of Post- and 
Rail-Launched Rockets 

The basic object of trajectory calculations for 
aircraft rockets is obviously to establish the relation 
of the position of the rocket at a given range to the 
position of the aircraft sight. The analytical expres- 
sion for the trajectory may contain three terms: 
(1) the gravity term, (2) the yaw term, and (3) the 
angular velocity term which is very small for firing 
from fixed launchers (either post or rail type) . The 
values of the terms as functions of propellant tem- 
perature, airspeed, dive angle, and slant range have 
been calculated for the various aircraft rockets and 
published in a number of reports. 1 

The trajectory drop (gravity term) is shown as 
the distance CE in Figure 5. It depends on: 

1. The rocket type, being smaller for higher 
velocity rockets; 

2. The propellant temperature, being smaller for 
higher temperatures because of the decreased burn- 
ing time; 

3. The dive angle, varying approximately as the 
cosine of this angle because of the different effective 
direction of gravity relative to the flight line; 

* See UBC reports listed in the CIT OEMsr-418 bibliog- 
raphy in the general bibliography in the appendix. 


/ 

4. The launching speed, decreasing with higher 
speed; and 

5. The slant range to the target. 

The yaw term consists of the product of the ini- 
tial yaw and a “launching factor.” When a rocket 
is fired into the airstream with an initial yaw to the 
stream, its trajectory is turned toward the direction 
of the relative wind by the action of the fins and the 
jet. The ratio of the angle through which the tra- 
jectory turns to the initial angle of yaw is called 
the launching factor / and may have values from 
1.00 down to less than 0.70, depending on the 
rocket type, the propellant temperature, the length 


f x (angle of affack of 
datum line + launcher 
angle) - launcher angle - 


f x (angle of attack of datum 
line + launcher angle L 



Launcher Line 


Datum Line and 
Zero Sight Line 

E f fee five 
Launching Line 
D Flight Line 


— Trajectorg Drop 

J ^L^ Tar 9 e ' h 


Figure 5. Factors in sighting for forward firing. 


of constrained motion on the launcher, and the 
indicated airspeed of the airplane. The evaluation 
of the initial yaw is complicated because the angle 
of attack (angle between flight line and boresight 
datum line [BSDL]) depends upon so many fac- 
tors — airspeed, dive angle, gasoline load, bomb 
load, etc. In addition, if there is a side wind, the 
plane axis will make an angle with the flight line 
in a horizontal plane, introducing horizontal as 
well as vertical yaws. Were it not that / is usually 
very close to 1.0, the problem of firing rockets from 
aircraft would be even more complicated. 

The angular velocity term is similar to the yaw 
term. If the rocket enters the airstream with an 
initial angular velocity, its trajectory is deflected 
in the direction of the angular velocity, and the ratio 
of the angle of deflection to the initial angular 
velocity is defined as the angular velocity factor. 
Its value varies from about 0.25 virtually zero. 

If angular velocity is negligible: 

Sight setting = trajectory drop + / X (angle of 
attack of datum line + launcher 
angle) — launcher angle 
= trajectory drop + / X angle of at- 
tack of datum line — (1 — /) X 
launcher angle. 


276 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


In case the launcher angle (i.e., the angle between 
the launcher and the zero sight line) is zero, only 
the first two terms in the sight-setting equation 
occur. For most aircraft, the fact that the sight is 
separated from the launchers by several feet adds 
the term h/R to the sight setting, where R is the 
slant range expressed in thousands of yards and h 
is the distance between the zero sight line and the 
launcher line expressed in yards. Sight settings are 
customarily expressed in mils. j 

24,6,2 Angle of Attack 

The most uncertain quantity in ordinary forward- 
firing problems is probably the angle of attack, the 
angle which the boresight datum line makes with 
the line of flight of the aircraft. It is necessary to 
differentiate between the true angle of attack and 
the effective angle of attack. The former is the angle 
between the BSDL and the undisturbed airflow at 
a great distance from the airplane. For simplicity 
in calculation of trajectories, it is customary to 
assume a uniform airstream around the airplane, 
although the direction of airflow adjacent to the 
airplane actually bears very little relation to the 
flight line, the effects of the propeller, fuselage, and 
especially the wings resulting in a flow which is 
uniform neither in magnitude nor direction. To 
circumvent this difficulty one defines the effective 
angle of attack to be that angle which gives the 
right answer in the sight-getting equation and then 
determines it experimentally for each aircraft under 
various flight conditions. The prediction of effec- 
tive angles of attack is an exceedingly difficult 
problem. For example, it was found that firing the 
5.0-in. aircraft rocket [AR] with and without the 
11.75-in. AR mounted in the airstream produced 
effective angles of attack differing by 10 mils, even 
after corrections for the differences in weight had 
been made. The problem is discussed in consider- 
able detail in Firing of Rockets from Aircraft, 1 * and 
an attempt at a theoretical understanding of it is 
made in reference 17 and in Exterior Ballistics. 1 

i At least two definitions of a mil are current. The standard 
Army mil is 1/6400 of a complete circle or 0.056250 degree, but 
in theoretical discussions it is more convenient to use the milli- 
radian, 0.057296 degree. The latter is 1.86 per cent larger 
than the former, but, for practical purposes when small angles 
are involved, either may be taken as 1-yd deflection in 1,000- 
yd range. 


24,6,3 Displacement and Drop 
Launchers 

The calculation of trajectories and sight settings 
for other launching methods involves no essentially 
new concepts and hence will not be discussed here. 
Because the initial conditions are more complicated, 
the sight-setting equations contain several more 
terms. The reader should consult Exterior Ballis- 
tics 1 Firing of Rockets from Aircraft, 1 * or refer- 
ence 18. 

24 7 RETRO FIRING 

Firing fin-stabilized rockets backwards from air- 
craft is no longer of much interest and will not be 
discussed here. Some ballistic calculations on the 
problem are given in reference 19. Firing fin-stabi- 
lized rockets accurately in any other direction is 
obviously impossible because of the large / factor. 

248 DISPERSION OF FIN-STABILIZED 
ROCKETS 

Dispersion is a measure of the scatter of the 
impact points of a group of identical rockets fired 
under supposedly identical conditions. Ordinarily 
this scatter is measured about the mean impact 
point of the group, but in some cases it may be 
measured about the point which one assumes would 
be the mean impact point of a very large group of 
rounds; e.g., the lateral dispersion may be meas- 
ured about the range line. Several different quanti- 
tative measures of dispersion are in current use, 
some of which are illustrated in Figure 6 (Figure 9 
of reference 20) . For our purpose we shall adopt the 
mean deviation as the measure of dispersion and 
shall use the two terms interchangeably. The mean 
deviation is computed by adding the deviations of 
the various rounds from the mean without regard 
to algebraic sign and dividing by the total number 
of rounds. Lateral dispersion is usually expressed 
in mils or in yards, and range dispersion in per cent 
of mean range or in yards. 

Many factors contribute to dispersion. Thus 
range dispersion may be introduced by variations 
in rocket weight, propellant weight, or effective gas 
velocity among different rounds of the group. Both 
range and lateral dispersion are affected by varia- 
tions in burning time, propellant temperature, or 


DISPERSION OF FIN-STABILIZED ROCKETS 


277 


wind velocity and by irregularities such as rough 
or crooked launchers, misaligned fins, and faulty 
lug bands. It was shown very early in the OSRD 
rocket developments, however, that finners fly quite 
straight after the cessation of burning and that the 


nozzle axis — and “gas malalignment” — the mal- 
alignment remaining when the mechanical malalign- 
ment is zero. No way is known to measure the gas 
malalignment directly, and it is usually inferred 
from an experimental test of dispersion by subtract- 



predominant cause of dispersion during burning is 
the malalignment. 

The malalignment of a rocket is usually defined 
as the distance between the center of mass of the 
rocket and the line of action of the thrust. Since 
the line of -thrust coincides approximately with the 
geometrical axis of the exit cone of the nozzle, a 
distinction is made between “mechanical malalign- 
ment” — the distance of the center of mass from the 


ing the effect of known mechanical malalignment. 
This procedure necessarily lumps together as gas 
malalignment all the other errors which can contrib- 
ute to dispersion, so that the result is always too 
large by an unknown amount. It is known, how- 
ever, that random variations of the thrust direction 
from the nozzle axis occur during burning, and the 
concept of gas malalignment is useful even though 
not precise . 


278 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


24 8 1 Dispersion of Finners in 
Ground Firing 

Whatever the cause or type of malalignment, its 
effect is a torque which causes the rocket to yaw 
and hence to deviate in the direction of that yaw. 
The resulting dispersion has been discussed quali- 
tatively in Chapter 21 and was first expressed quan- 
titatively in The Effect of Fin Size, Burning Time, 
and Projector Length on the Accuracy of Rockets, 21 
a report which has become a classic in rocket liter- 
ature^ In this analysis, damping, drag, gravita- 
tional force, and cross-wind force are assumed negli- 
gible, and the equations of motion of the projec- 
tile are solved assuming the malalignment torque 
to be constant and the restoring torque of the fins 
to be proportional to the yaw and to the square of 
the velocity. The solutions of the equations turn 
out to be Fresnel integrals, and they are plotted in 
Figure 7 from which the qualitative conclusions 
listed in Chapter 21 and many others, can be de- 
duced. If one considers a particular projectile, the 
ordinates in Figure 7 are proportional to deflection 
of the trajectory in the plane of the yaw per unit 
malalignment, the abscissas are proportional to 
time, and the parameter is essentially projector 
length. Thus each curve shows the variation in 
trajectory deviation (the angle 0 in Figure 1) with 
time during burning, and, since the rocket, after 
burning continues its flight in the direction it was 
pointing when the thrust ceased, putting the value 
tb into the graph gives the trajectory direction at all 
times after burning. To convert this to lateral devi- 
ation of the impact point, which is most frequently 
of interest, one must multiply by the sine of the 
angle between the plane of yaw and the vertical 
plane and divide by the cosine of the angle of eleva- 
tion of the launcher, the small correction for the 
burning distance usually being neglected. It will 
be noted that two abscissa scales are included, the 
upper one applying at all times and the lower one 
being appropriate only at the end of burning. 

Despite the seemingly rather restrictive assump- 
tions on which the theory is based, it has been found 
to be in excellent agreement with experiment. The 
fact that the malalignment torque is not constant 
either in magnitude or direction during burning 
does not invalidate the conclusions with regard to 

k There are several earlier CIT reports on the same subject. 
A later one 22 includes simplified formulas useful in restricted 
regions. 


variation of dispersion with launcher length, burn- 
ing time, or fin size (i.e., the yaw oscillation dis- 
tance a, defined in Section 24.5 and in Chapter 21). 
In particular it is important that the theory holds 
fairly well even for supersonic velocities where the 
restoring moment is not proportional to the square 
of the velocity, because normally a rocket reaches 
the flat portion of the curve before the square law 
breaks down so that practically all of its deflection 
is acquired in the low-velocity region where the 
theory is valid. 

The chief limitations of the theory are the diffi- 
culty of determining the effective launcher length p 
(p stands for “projector”) and the actual malalign- 
ment L 0 during flight. An accurate definition of p 
would be the distance through which the rocket is 
constrained to move with zero deflection, but 
whether this constraint ceases when the head or 
front lug leaves the launcher or continues as long 
as the center of mass, the tail, or the rear lug is 
in contact is seldom obvious a priori. Analysis of 
a large number of firings of the 4.5-in. barrage 
rocket, for example, showed that the data could 
best be brought into agreement with the theory by 
assuming that the constraint ceases when the tail 
leaves the launcher. 23 Probably this is approxi- 
mately the case for most relatively lightweight 
rockets on rail or tubular launchers. 

On the other hand, the 5.0-in. HVAR was found 
to be about equally accurate from zero-length 
post launchers or the 73^-ft Mk 4 launcher even 
in ground firing. The curves of Figure 7 provide 
the explanation 24 of the behavior of the HVAR. 
This rocket has an initial acceleration of 55gr, from 
which it is easily calculated that its rear lug will 
clear a 7.5-ft launcher in approximately 0.09 sec- 
ond. Since its effective burning time at 70 F is 0.9 
second and its yaw oscillation distance a is 320ft, the 
values of t\/ Q/g- which we need for the graph are 
2.1 at the end of burning and 0.21 at 0.09 second, 
and the ordinates of the curve p/a = 0 correspond- 
ing to these times are respectively 0.010 and 0.0006. 
Thus, if fired from a zero-length launcher, this 
rocket will acquire 0.06 of its total deflection during 
burning in the first 7.5 ft. If its average deflection 
at the end of burning is 20 mils, then at 7.5 ft it will 
be 1.2 mils. Since the separation between the two 
suspension lugs is approximately 36 in., they will 
undergo a relative lateral displacement of 0.043 in. 
in the first 7.5 ft. But the clearance between the lugs 
and the slot of a Mk 4 launcher is approximately 0.060 


DISPERSION OF FIN-STABILIZED ROCKETS 


279 


.on 


B 


.010 


.009 


6 C - Deflection (mils) 

L 0 = Malalignment ( aoo / in.) 

H = Radius of gyration of rochet (ft) 
Effective projector Length 
G = Acceleration of rochet ( ' ft/sec V 
v = Gt~ Velocity (ft/sed 
v b = Velocity at end of burningffl/sec) 
t •lime from start of burning (sec) 
t b . Time of burning (sec) 

T - Period of yaw oscillation at 
< v b (sec) 

v b T= "Wavelength* of yaw 


p/jr 




r- 


oscillation (ft) 


Asymptotic Values 

p/o-'Q 


J< .§c 
CT L a 





0.04 0.16 0.36 Q64 1.00 1.44 1.96 ZS6 324 4,00 

For 


Figure 7. Deflection of trajectory by malalignment (zero launcher velocity). 


in. Hence it is clear that it subjects the average 
round to little or no constraint and its effective 
length must be almost zero. Even if the clearances 
were made quite small, it is doubtful that the dis- 
persion of a rocket as heavy as the HVAR would 
be improved, because it is not feasible to build an 
aircraft launcher of sufficient weight and rigidity 
to constrain it effectively. 

More detailed applications of the theory to vari- 
ous CIT rockets are given in references 25 and 2G, 
and one more will be included here. The chemical 
warfare rocket [CWR] (see Section 18.4), before the 


addition of radial fins, had the following charac- 
teristics: 

Velocity at the end of burning (70 F): F fc = 710 fps; 
Yaw oscillation distance: <7 = 280 ft; 

Radius of gyration (see Table 2 in Section 21.3): 
K =1.22 ft. 

Hence 

K 2 

— = 0.0053; 

G 

_ 2.54. 
a 


o 


280 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


Assuming an effective launcher length p = 8.4 ft and Using this value and 0/Lo = O.9, we find for the dis- 
a burning time ^ = 0.51 second, we have persion expected at 45 degrees elevation angle 


£ = 0.03; 

a 


0.9 X 25 X 2 
7 r cos 45° 


20.3 mils. 


V btb 


= 1.3. 


From Figure 7 1 we read 


K 2 0 0 

— • j- = 0.0053; j- « 0.0048. 

<7 Li o Lq 


Hence 


y- « 0.9 mils per 0.001-in. malalignment. 
T o 


Increasing the burning time above this value would 
make no significant change in dispersion, since all 
values of (. K 2 /a)(d/L 0 ) for longer burning times 
fall between 0.0046 and 0.0053. Reducing the burn- 
ing time by half, on the other hand, giving 


results in 


so that 


V btb 


= 0.64 


K 2 ' B_ 
<j L o 


0.0026, 


Y~ — 0.49 mils per 0.001-in. malalignment. 

Lq 


It is easily seen also from the graph that the 
same improvement in dispersion without reduction 
in burning time could be obtained with p/ a = 0.09, 
that is, triple the original effective launcher length, 
if such a launcher were practicable. 

The theory is primarily useful for making com- 
parisons of this type, but it can be used to predict 
the actual magnitude of dispersion if something 
is known about the average malalignment to be 
expected in practice. From measurements on vari- 
ous rockets it is known that the minimum attainable 
gas malalignment is approximately 1 mil and values 
of 2 or 3 mils are common. Since the CWR has its 
center of mass 25 in. ahead of the nozzle throat, 
1 mil corresponds to a malalignment of 0.025 in. 

1 The subscript c on the 0’s in Figure 7 signifies merely that 

it was calculated on the assumption of constant acceleration 
and constant malalignment. 


The factor 2/tt occurs because the directions of the 
malalignments are randomly distributed. One 
would expect 20 mils to be an extremely optimistic 
estimate of dispersion, and in fact the actual dis- 
persion is nearer 45 mils, indicating that 2 mils 
would have been a better guess at the gas malalign- 
ment. If the rockets are not well made, the mechan- 
ical malalignment may further increase the disper- 
sion, but with careful manufacturing and inspection 
it is usually possible to keep the mechanical mal- 
alignment small enough so that its effect is com- 
pletely obscured by the gas malalignment. 

The theory indicates the following possible ways 
to increase the accuracy of a fin-stabilized rocket: 

1. Increase the launcher length; when it can be 
done, this will always reduce the dispersion if the 
rocket is actually constrained by the launcher, but 
it is seldom practicable. 

2. Decrease the burning time; this will be effec- 
tive only if it can be decreased to the point where 
it is considerably less than the period of yaw oscil- 
lation <r, which is seldom possible for high-perform- 
ance rockets. 

3. Reduce < 7 ; i.e., increase the stability by using 
larger fins or (usually less feasible) by moving the 
center of mass forward. The difficulty here is that 
rather large increases in fin size are required to 
affect o- appreciably (see Chapter 19 for the effect 
of fin size on the HVAR), and these are usually 
precluded by space considerations. 

4. Increase the radius of gyration K ; in other 
words, design a new rocket that is longer and 
slimmer. 

5. Reduce the gas malalignment. 

Much effort has been expended in attempts to 
reduce gas malalignment by some variation in the 
interior of rocket motors. A number of the expedi- 
ents tried are discussed in reference 27. None 
showed any promise of success. Apparently gas 
malalignment is rather fundamentally tied in with 
high-speed gas flow and cannot be significantly 
reduced. Its effect can be partially circumvented, 
however, by two expedients: using multiple nozzles 
and rotating the rounds. Apparently the directions 
of gas malalignment in various nozzles of one plate 
are at least partially independent and tend to cancel 


DISPERSION OF FIN-STABILIZED ROCKETS 


281 




Figure 8. Deflection of trajectory by malalignment in aircraft forward firing from “zero-length” launchers. 




282 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


one another out. Thus the gas malalignment of the 
eight-nozzle HVAR is less than 0.87 mils. 28 Except 
for the very atypical target rocket, no serious 
attempts to reduce dispersion by rotating fin-sta- 
bilized rockets were made by CIT. The British and 
the Section H 29 workers have tried it with some 
success, however, and the theoretical possibilities 
along this line are discussed briefly in Exterior 
Ballistics . lc 

24,8,2 Dispersion in Firing Finners 
Forward from Airplanes 

By a simple change of variable, the curves in 
Figure 7 can be adapted to the case where the 
rockets have a relative velocity with respect to the 
air at the beginning of burning, thus giving the dis- 
persion caused by malalignment in forward firing 
from aircraft. This theory is derived in reference 
30 and reduced to graphical form in reference 31. 
Since one more parameter, the airspeed of the plane, 
now appears in the theory, it is not possible to show 
the whole story on a single graph, and only the 
curves for the most important case, zero launcher 
length, are reproduced here as Figure 8. An exami- 
nation of the complete set of curves leads to the fol- 
lowing conclusions: 

1. For given values of p/<r and V a /Vb (aircraft 
velocity divided by rocket burnt velocity relative 
to the aircraft) , the dispersion increases rapidly with 
burning distance d b , reaching a maximum at a value 
of d b /a- between 0.2 and 0.5 and then decreases 
(except for extremely small V a ) for longer burning 
distances. 

2. Dispersion decreases with increased launcher 
length, but this effect becomes less marked at higher 
airplane speeds. Thus for the HVAR (V b = 1,350, 
d b « 600, a « 300), theory predicts the following 
decreases in dispersion in going from a zero-length 
to a 6-ft launcher: 

Ground firing 44 per cent; 

Airspeed 270 fps 41 per cent; 

Airspeed 540 fps 36 per cent. 

In practice, the improvement is likely to be con- 
siderably less than this, as pointed out in the pre- 
vious section. 

3 . The relative gain in accuracy when going from 
a stationary to a moving launcher is greatest for 
the zero-length launcher, where it can amount to a 
factor of 10 or more. The burning distances of CIT 


aircraft rockets are all so long as to place them well 
beyond the maxima in Figure 8. In this case, 
asymptotic formulas are applicable, and the 
single convenient curve of Figure 9 (taken from 
reference 32) covers all cases. 

The dispersion of the ammunition itself is by no 
means the predominant effect in forward firing, 
however. Many other factors contribute to the 
inaccuracy, such as sighting errors, faulty estima- 
tion of range to the target, airspeed and dive angle 
incorrect for the sight settings used, uncertainty 
in the temperature of the ammunition, random 
winds, and firing while the dive angle is changing. 
The quantitative effects of these various errors are 
analyzed in reference 33 , from which is taken Table 
3, showing a typical case. Many good illustrations 
of these effects are given in reference 16. 

Table 3. Effect of various factors on dispersion in 

forward firing of 3.5-in. AR from TBF-1. 


Launcher 3° above datum line Dive angle 20° 

Mean temperature 70°F Range 750 yd 

Launcher length 7.5 ft 


Vertical dispersion (mils) 
Aircraft speed (knots) 200 225 

250 

275 

Ammunition dispersion 

5 4 

3 

2 

Pure aiming error 

3 3 

3 

3 

Random wind (10 fps) 

.3 3 

3 

3 

Range error (75 yd) 

2 2 

2 

2 

Temperature error (10°F) . . . . 

2 2 

2 

2 

Aircraft speed error (10%) . . . 

6 6 

6 

6 

Total 

9 9 

8 

7 

Lateral dispersion (mils) 
Ammunition dispersion 5 4 

3 

2 

Pure aiming error 

2 2 

2 

2 

Wind (10 fps) 

. . 8 8 

8 

8 

Total 

.10 9 

9 

8 


249 UNDERWATER TRAJECTORIES 

The underwater ballistics of fin-stabilized rockets 
has already been briefly introduced in Chapter 15 
in connection with head shapes. We have seen that 
the projectile after entering the water travels in a 
bubble and is in contact with the water only near 
the nose and the tail. In this position it effectively 
has a yaw with its trajectory; consequently the 
forces of the water reacting on the nose are not in 
general symmetrical, and a net cross force exists on 
the nose. In the case of a pointed projectile this 
cross force is in the direction opposite to the side 
of the bubble on which the tail lies, and hence is 


UNDERWATER TRAJECTORIES 


283 


usually an upward force because both the effect of 
gravity and the initial impulse of the water on the 
nose tend to make the tail ride on the bottom of 
the bubble. It has been demonstrated that the 
amount of this cross force varies greatly with the 
shape of the ogive. Thus there should be practically 
no side force on a hemispherical ogive, since it pre- 
sents the same form to the water when rotated 


Two other factors in addition to nose shape deter- 
mine the magnitude of the cross force. The length 
is important because a short rocket will have to 
have a larger yaw in order to ride on the bottom of 
the bubble than will a longer rocket of the same 
diameter and head contour. Thus an ordinary shell, 
whether spinning or not, is so short that it cannot 
be stabilized under water at all, but turns sideways 



Figure 9. Dispersion of long-burning rockets in forward firing. Middle curve is exact for very long-burning 
rockets; side curves illustrate approximate formulas. 


through a small angle. There will still be a side 
force on the tail, but, for a long slender projectile 
in which the center of gravity is near the ogive, this 
should be negligible. On the other hand, for sharper 
ogives the side force is greatly increased, since for 
a given yaw the ratio of the amount of water forced 
to one side to the amount forced to the other side 
of the projectile is greater the longer the ogive. 


and comes to rest almost immediately. The effect 
of a motor of diameter less than that of the head is 
to increase the yaw, because the smaller motor 
must dip farther into the side of the bubble in 
order to acquire a given restoring moment. Thus 
for the 5.0-in. aircraft rocket which has a 3.25-in. 
motor, no head shape was found which would make 
the rocket stable under water. 



284 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 


If the side force becomes too great, as it may at 
high entry velocities and large entrance angles, the 
rocket breaks in two, usually at the junction be- 
tween the motor and the head, and the head is 
brought to rest almost immediately. Otherwise the 
side force produces a curvature of the trajectory, 
and it is easily shown that the path approximates 
an arc of a circle, the radius of which is directly 
proportional to the rocket’s mass and inversely 
proportional to the cross force. 

If this simple picture were always exactly repro- 
duced in practice, every rocket would follow an 
upward-curving path and have a trajectory as 
shown in Figure 10 until its velocity was so reduced 
that gravitational forces became appreciable. If 



Figure 10. Ideal underwater trajectory of a fin- 
stabilized rocket, assuming negligible change in 
velocity while below the surface. 

fired so as to enter the water at a sufficiently small 
angle with the surface, it would emerge making the 
same angle, and the horizontal distance between 
the entrance and exit points would be proportional 
to the sine of the entrance angle. Although the 
limits of error are necessarily rather large, the 
experimental firings indicate that the average rocket 
does have such a trajectory. Also in accordance 
with the theory, it has been found possible to con- 
trol the radius of curvature within the limits where 
the rocket can stand the cross force and to reduce 
the deceleration coefficient substantially by shaping 
the heads so that the water breaks away from them 
at a smaller diameter and forms a smaller bubble 
as discussed in Chapter 15. 

Nevertheless, very erratic behavior is exhibited 
by a small percentage of the rounds, and little is 
known about the reasons for it. One would surmise 
that a yaw at the instant of water impact might 
throw the rocket to one side of the bubble and thus 
cause the normal curvature of the trajectory to take 
place in a plane inclined to the vertical. British 


firings under conditions which allowed recovery of 
the rounds showed that motor tubes (with thinner 
wall than American designs) sometimes become dis- 
torted by the impact forces and that occasionally 
one of the four fins remains on the motor; in either 
of these cases a steering action on the rocket results. 
A bizarre example of what kinds of things may hap- 
pen was provided by a Tiny Tim which ricocheted 
apparently normally and landed on shore, but when 
recovered was found to have a 1-ft length from the 
front of the motor tube missing, the head being 
jammed back into the remaining tube and in fairly 
good alignment. 

24.9.1 Tactical Effectiveness of 
Underwater Rockets 

The ability to vary both the curvature of the 
trajectory and the rate of loss of velocity under 
water makes possible a significant increase in the 
effectiveness of rockets with certain head shapes 
under certain conditions. A brief quantitative dis- 
cussion of this point is contained in reference 34 
from which all of the theory of underwater trajec- 
tories has been taken. Additional theory is dis- 
cussed in references 35 and 36. Qualitatively, it is 
evident that the curvature of the trajectory under 
water causes a deflection from the straight-line air 
trajectory, which in certain cases may send the 
rocket into the target, thus increasing the proba- 
bility of a hit, but in other cases may send it away 
from the target, decreasing the probability. The 
rapid deceleration of the rocket under water causes 
it rapidly to drop below a velocity at which it can 
cause significant damage, and this factor, as well 
as the curvature of the trajectory, must be evalu- 
ated to determine the rocket’s effectiveness under 
various conditions. For example, consider the case 
of a submerged submarine, represented in Figure 11 
by the circle GHI where the water surface is DEF. 
ADG and CFI are the extreme trajectories, having 
an entrance angle a, that just reach the target. The 
plane MN is perpendicular to the air part of the 
trajectory. The effective target area then extends 
from J to L and is significantly wider than the actual 
target, if the underwater path FI is short enough 
so that the rocket reached I with a velocity great 
enough to cause significant damage. If, however, 
the velocity at I is below that specified to produce 
the desired damage, a third trajectory must be laid 


UNDERGROUND TRAJECTORIES 


285 


out such that the underwater part of it is equal to 
the length of underwater travel required to bring 
the rocket down to the limiting velocity for damage, 
and the effective target area (proportional to the 
distance between AD and the air part of this new 
trajectory) will be correspondingly reduced. The 
interrelation of these various factors makes the 
choice of the optimum head shape and entrance 
angle a rather difficult one, depending very criti- 
cally on the type of target. 




1 





Figure 11. Effective target area for submerged 

cylindrical target. 

24 10 UNDERGROUND TRAJECTORIES 

Firings of 5.0-in. HVAR’s and 11.75-in. AR’s 
into earth have provided additional verification of 
the theory of underwater trajectories, since one 
would expect underground and underwater per- 
formance to be qualitatively similar. That a rocket 
travels under ground in a “bubble” is apparent from 
the erosion marks exhibited by recovered rounds 
(see Figures 13 and 14). Thus heads which have 
long straight underwater trajectories (small cross 
force) actually do give superior performance under 
ground. 

Because of the variable consistency of earth and 
the meagerness of the data, it is difficult to make 
any general statements about underground trajec- 
tories other than that the much larger forces require 
heads giving less nose lift than is usable under water. 
Heads with little or no lift may still be unsatisfac- 
tory, however, if their drag is large so that the axial 
force on the motor is increased. For any head shape, 
it is essential that the motor tube have a relatively 
thick wall and that its junction with the head be 
strong. 



Figure 12. HVAR head shapes tested for under- 
ground trajectory. 


CIT tests of ground penetration of the 5.0-in. 
HVAR are discussed only in the weekly progress 
reports. 37-39 The head shapes tested are shown in 
Figure 12 and the results are summarized as follows: 


286 


EXTERIOR BALLISTICS OF FIN-STABILIZED ROCKETS 



Figure 13. 11.75-in. sphere-ogive head after 

rocket had penetrated 75 ft of earth with impact 
velocity 1,275 fps. ‘‘Wart” on nose is fuse. Note 
that erosion extends only to intersection of sphere 
with 20-caliber portion (O^-in. diameter). 


Type 1: Standard underwater sphere-ogive head. 
With this head, the rocket was entirely stable, 
traveling underground (in clay covered with sand) 
an average of nearly 50 ft, at dive angles of 15 to 20 
degrees. The lateral deviation was very small, but 
deviations averaged several degrees from the mean, 
and, in contrast to underwater results, a consider- 
able upward curvature of the trajectory was noted. 
The average round was recovered at a depth slightly 
less than half that corresponding to an extension of 
its air trajectory, and one, fired at 15-degree dive 
angle, actually emerged after 24 ft underground and 
detonated in the air (it carried a deceleration-dis- 
criminating fuze) . 

Type 2: Standard semi-armor-piercing Mk 2 
head. All heads separated from their motors and 
emerged after 6 to 8 ft of travel underground, show- 
ing erosion on the nose and one side. 

Types 3 and 4: Modifications of Type 2. Per- 
formance identical with Type 2. Apparently a por- 
tion of the ogive back of the spherical nose remained 



Figure 14. Special 11.75-in. AP head after rocket had penetrated 65 ft of earth with impact velocity 1,225 
fps. Note that erosion extends to full 11.75 in. diameter. 


UNDERGROUND TRAJECTORIES 


287 


in contact with the “bubble,” giving a large cross 
force. 

Type 5: Blunt ogive head. Performance identical 
with Type 2. 

Type 6: Hemispherical-nose head. All heads 
broke off from their motors, but the erosion was 
more symmetrical than in the unstable cases pre- 
viously mentioned, indicating that drag rather than 
cross force may have been the primary factor in the 
failure . 

Type 7: Sphere-cone ogive. This head appeared 
to be near the limit of stability since, although all 
heads broke off, they penetrated 12 to 14 ft and 
eroded quite symmetrically. Apparently the drag 
with this size of spherical nose is still too great. 

Earth penetration tests with Tiny Tim have been 
made by NOTS, Inyokern, and one must consult 
Navy reports for the details. One such test gave 
the following results at impact angles of approxi- 
mately 33 degrees: 


Sphere-ogive head (Figure 13). This penetrated 
70 ft in the same direction as the air trajectory for 
1,275-fps striking velocity. 

Mk 1 head. For striking velocity of 1,380 fps, 
penetration averaged 50 ft and the rounds turned 
up 10 degrees from their air trajectory. (At shal- 
lower angles one round broke.) 

Special heavy head (Figure 14) having same ex- 
terior contour as Mk 1 but a greater length and 
weight. These rounds weighed approximately 1,550 
lb instead of 1,120 as for the previous types. For an 
entrance velocity of 1,225 fps, their penetration 
characteristics were identical with those of the 
sphere-ogive heads. Why these should not turn up 
as the Mk 1 heads do has not been explained. 

When heads having the same shape as the Mk 1 
were fired with Mk 2 motors (wall thickness 0.240 
in. instead of 0.300 in.), all motors were shattered, 
although their underwater performance is entirely 
satisfactory. 


Chapter 25 

EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 

By C. W. Snyder 


25 1 SIMPLEST TYPE OF 

SPINNER MOTION: NUTATION 

T he motion of a spin-stabilized rocket in the 
absence of gravitational and aerodynamic forces 
is closely analogous to that of a finner in air. For 
the latter, the equilibrium position is one of zero 
yaw, and if displaced from it the rocket oscillates 
(in the plane determined by its axis and the tangent 
to the trajectory) with a frequency which increases 
with velocity at just the proper rate so that the 
distance traveled in each oscillation is a constant, a. 
The equilibrium position of a spinner in the absence 
of air is also one of zero yaw. When displaced from 
this orientation, it oscillates so that the distance 
traversed during each oscillation cycle is a constant, 
X, analogous to a. Unlike the finner, however, the 
oscillations are not in one plane — the nose of the 
rocket moves in a spiral about the trajectory of 
the center of mass. This motion is called nuta- 
tion; its projection on a plane through the trajec- 
tory duplicates exactly the oscillation curve of a 
finner. The constancy of the distance covered in 
each nutation cycle is a consequence of the fact that 
the rate of nutation is proportional to the rate of 
spin, which is, as indicated in Chapter 21, propor- 
tional during burning to the velocity. The analogy 
between finner and spinner motion is exact both 
during and after burning if one assumes that there 
is no jet malalignment, no aerodynamic forces on 
the spinner, and no damping forces on the finner. 

Although these features of similarity between 
spinner and finner behavior are helpful, both the 
force system and the motion of spinners under con- 
ditions of reality are, in general, somewhat more 
complicated than those of finners. The complica- 
tions result from the larger number of forces and 
moments which act on spinners, in combination 
with gyroscopic action. 

25 2 FORCE SYSTEM OF SPINNERS 

As was done for finners in Chapter 24, the first 
step is to set up a system of a small number of forces 


and torques which will be equivalent in effects to 
the multiplicity of distributed forces, both internal 
and external, which govern the motion of spinners. 
A detailed discussion of such a force system is given 
in Exterior Ballistics. 1 

The important elements of the system are five 
forces and four moments, as tabulated below. 

The forces are 

1. Gravity. 

2. Jet forces, which act only during burning. 

3. The drag , which, like that for finners, results 
from high air pressure on the nose, reduced pressure 
behind the rocket and skin friction. 

4. The lift or cross-wind force , which accompanies 
yaw and causes planing action, tending to push the 
rocket in the direction of its yaw. 

5. The Magnus force, an aerodynamic force pecu- 
liar to spinning projectiles. a It appears whenever 
there is a component of airflow perpendicular to the 
spin axis (i.e., when the yaw is not zero) and tends 
to move the rocket in a direction perpendicular to 
both the yaw and the trajectory. We can visualize it 
most easily if we consider the case where the rocket 
is oriented broadside to the relative wind (yaw = 
90 degrees). The skin friction carries a certain 
amount of air around with the rocket as it rotates, 
and on one side of the rocket this trapped air col- 
lides wdth the air flowing past, creating a higher 
pressure, while on the other side the trapped air and 
the free air flow in the same direction giving reduced 
pressure. Theoretical analysis shows that the Mag- 
nus force is proportional to the product of the 
rocket’s angular velocity by its linear velocity, and 
for smaller yaws than 90 degrees the factor sin 8 is 
also included. 

The most important moments are: 

1. The overturning moment, which tends to turn 
the rocket across the trajectory because the center 
of pressure (where lift and drag are assumed to act) 
lies forward of the center of mass. Finners have a 
righting moment instead. 

2. The Magnus moment exists whenever the 

a This is the force which causes a properly thrown baseball 
to curve. 


288 


MOTION DURING BURNING 


289 


Magnus force is applied elsewhere than at the 
center of mass. It is small in magnitude but im- 
portant in effect. 

3. The spin deceleration moment, which tends to 
slow down the spin because of air friction. 

4. The damping moment, which always opposes 
the yaw, exists only when the yaw is changing and 
tends gradually to damp it out. It results from the 
difference in the forces on the two ends of the 
rockets associated with their different air velocities 
when the yaw is changing. 

The greater complexity of these forces and mo- 
ments as compared with those which act on a fin- 
stabilized projectile is apparent. For a finner, force 
5 and moments 2 and 3 are entirely absent, while 
force 4 averages to zero because the rocket has a 
zero yaw on the average. If it is assumed that the 
overturning moment is proportional to the yaw 
angle (as was done also for finners and is approxi- 
mately true for small yaws), then the equations of 
motion are linear, and the effects of the various 
forces may be computed separately and added to 
give the final motion. We shall confine ourselves 
mainly to this approximation since it will explain 
adequately the main features of spinner motion. 
There remain, however, a few important effects 
that require more complicated analysis. 

The general features of spinner motion were 
sketched in Chapter 21, and it is suggested that the 
reader glance through the pertinent sections there 
before proceeding further. In the following para- 
graphs, we shall extend the analysis of Chapter 21, 
but without discussing the equations of motion 
from which the results are calculated. For further 
details the reader is referred to Exterior Ballistics 1 
or to the original papers 

It will clarify the following to keep in mind a 
particular rocket, and the 5.0-in./5 HCSR Model 
34 (5.0-in. Rocket Mk 10 Mod 0) will serve as an 
example. It is described in Chapter 20. In Table 1 
are given the pertinent ballistic constants for such 
a rocket. Slight changes from the actual constants 
have been made for convenience in applying the 
graphs to follow. Notation used in this chapter is 
the same as in Chapter 21, with certain additions, 
and is summarized in Table 2. 

25 3 MOTION DURING BURNING 

As indicated in Section 25.1, a spinner, in the 
absence of gravity and aerodynamic forces, will 


move along a straight trajectory with its nose oscil- 
lating in a spiral of constant nutation distance. In 
a real rocket, of course, this motion is modified. 
During the period of propulsion (burning period) 
the principal factors affecting the motion are the 
overturning moment, gravity, interaction with the 
launcher, and wind. 

Table 1 . Ballistic constants of typical 5.0-in. spinner.* 


Stability factor during burning: S = 2 
Radii of gyration: K 2 = 0.60 ft 2 

k 2 = 0.030 ft 2 

j = V20 « 4.5 


Feet per turn: 
Feet per nutation: 
Burning distance: 


v = 6 ft 

X = 120 ft 
d b = 325 ft 


Velocity parameter for the end of burning: 

r _ \dl _ / 325 _ ! 

Acceleration at 70 F: G — 30 g = 966 ft/sec 2 

= V2)^/G = J— = 0.50 sec 
\ 966 


50 


F\ = V2 GX = V240 X 966 = 481 ft/sec 


1 

^X 


0.00218 sec/ft. 


* The constants tabulated are approximately those of the 5.0-in./5 HCSR 
Model 34 which has an overall length of 32 in. (including nose fuze), a 
weight of 50 lb, and a velocity of 790 fps, and spins at 130 rps. 


25.3.1 Eff ec t 0 f Overturning Moment 

The overturning moment, the principal aero- 
dynamic effect, introduces gyroscopic precession. 
Any uniform torque on a spinning gyroscope causes 
its axis to precess so that the motion of any point 
on the axis is a circle. The overturning moment 
acting on a spinner with a given yaw leaves the mag- 
nitude of the yaw constant but rotates the plane of 
yaw uniformly about the trajectory. In general, 
the initial conditions are not such as to give this 
dynamically stable mode of motion, but the nuta- 
tions will be superimposed on it. 

In the following discussion we shall frequently 
find it convenient to represent spinner motion 


290 


EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 


Table 2. Notation for spin-stabilized rockets. 


f = velocity parameter, f = Vrf/X; = Vp/X. 

0 ss characteristic function for a trajectory orientation 
(see Table 3). 

0 = orientation of the tangent to the trajectory relative to 

the launcher. 

do = quadrant elevation of the launcher. 

X = distance traveled in one nutation, assuming con- 
stant S (ft) . 

v = distance traveled in one rotation. 

4> = characteristic function for orientation of rocket axis 
(see Table 3) . 

<t> = orientation of the rocket axis relative to the launcher. 
b = subscript denoting “at the end of burning.” 
d = distance along trajectory from point of ignition (ft). 
d = hGt. 

E = function giving variation of malalignment effect with 
launcher length. 

G = acceleration of the rocket in horizontal fire (ft /sec 2 ). 
g = acceleration of gravity (32.2 ft/sec 2 ). 

K = transverse radius of gyration (ft). 
k = polar radius of gyration (ft) . 

1 = length of rocket (ft). 

p = launcher length (ft). As a subscript, it signifies “at 
the end of the launcher.” 

q = transverse angular velocity of mallaunching (radians 
per second). As a subscript, it denotes “produced by 
mallaunching.” 

R m = jet malalignment (ft) . 

S = stability factor [see equation (23) of Chapter 21]. 
s s= spin angular velocity (radians per second). 
t = time (seconds). 

t\ = time required to complete first nutation (assuming 
nutation and acceleration to commence simultaneously 
and rocket to continue burning throughout the nuta- 
tion). tx = V2X/G. 

u = unbalance. Subscripts S and D denote static and 
dynamic unbalance. 
v = velocity (fps). 

v\ = velocity at the end of the first nutation (same assump- 
tions as for t\). V\ = V2G\. 

W = wind velocity (fps). As a subscript, it denotes “pro- 
duced by wind.” 

Wn ^ wind velocity component perpendicular to launcher 
(fps).. 

X, Y = coordinates in a plane perpendicular to the launcher. 
X is positive to the right and Y is positive down. 


graphically by using a moving system of coordinates 
having its origin at the center of mass of the rocket, 
its Z axis pointing in the direction of the launcher, 
its X axis pointing to the right, and its Y axis 
pointing down. The change of the rocket from its 
original orientation (the Z axis) can then be repre- 
sented by the projection on the XY plane of a point 

1 ft ahead of the center of mass, and, in the approx- 

imation of small angles, the distance of the pro- 
jected point from the origin is proportional to the 
orientation angle. As the motion proceeds, this 
point will trace out a curve which is easily inter- 


preted by imagining one’s self standing behind the 
rocket and watching the motion of the nose. Such 
curves we shall call “orientation curves,” and a 
number of them will be included later in this chap- 
ter. A much more complete set is contained in 
Exterior Ballistics. * 1 

As in the previous chapter, we shall use three 
angles to specify the rocket’s position and motion: 
6 = angle between the launcher line and the tan- 
gent to the trajectory, 

<f> = angle between the launcher line and the rocket 
axis, and 

8 = angle between the rocket axis and the tangent 
to the trajectory . b 

Since the motion is not plane, we shall have to give 
the projections of these angles on the horizontal 
and vertical planes, and shall denote the projec- 
tions by subscripts X and Y, respectively. 

The orientation curve for a precession or an un- 
damped nutation is a circle, and it is simple to super- 
impose the two circular motions provided that we 
know their relative velocities. From an analysis 
which includes the effect of the overturning moment, 
but excludes other aerodynamic forces and gravity 
(which would introduce only minor corrections) , we 
find 

Angular velocity of nutation = 

qJU2 

2^(1 + Vl — 1/>S); 

Angular velocity of precession = 

|£(1 - Vl - l/S). 

From the ratio of these we find that the number 
of nutations for each precession is 

1.00 for S = 1.00 (very low stability factor); 

5.82 for S = 2.00; 

9.86 for S = 3.00; 

4*8 — 2 as the limit approached for very large S. 

Thus the distinction between nutations and pre- 
cessions virtually disappears for very low stability 
factors. 

At the same time, of course, the rocket is rotating 
(spinning) about its oscillating axis with a higher 
angular velocity s. Dividing this by the angular 


b Evidently in order to draw an orientation curve for the 
yaw angle 5, we should have to take the Z axis pointed along 
the trajectory instead of along the launcher, but only one 
such curve is given in this book. 


MOTION DURING BURNING 


291 


velocity of nutation, we get the number of spin 
rotations per nutation as 

2.0 tv for S = 1.0; 

At 

1.17 ~ for S = 2.0; 

1.10 ~ for S = 3.0; 

K 2 

1.00 as the limit approached for large S. 

Since v is the distance traveled during each rota- 
tion, the distance for each nutation is, for large 
values of S, 

. _ vK 2 
X k 2 ' 

This depends only on geometrical constants of the 
rocket. For values of S customarily used for 
ground-fired spinners, this expression gives a result 
about 15 per cent lower than that observed. 


0 6 



Figure 1. Precession and nutation without damp- 
ing (S = 2). 


The orientation of a precessing and nutating 
rocket with a stability factor of 2 is shown graphi- 
cally in Figures 1 and 2. The first shows the case 
where the nutation amplitude is constant and one- 
fourth that of the precession, and the second shows 
a case of extremely large damping where the ampli- 


tude of the nutation decreases to 0.7 times its 
former value during each nutation and where the 
rocket is released with zero yaw, so that initially 
the nutation and precession amplitudes are equal. 
The numbers along the curves indicate the ends of 
each nutation. 

3 



Figure 2. Precession and damped nutation (S 
= 2 ). 


25 3 2 Effect of Gravity 

If no aerodynamic forces were acting, the effect 
of gravity would be simply a vertical drop of the 
trajectory. Thus our hypothetical HCSR fired 
horizontally from a zero-length launcher would 
have an acceleration g downward and 30 g forward 
so that its center of mass would move in a straight 
line falling below the horizontal by an angle whose 
tangent is 1/30, i.e., by 33 mils. Since its nose 
would continue to point in the direction of launch- 
ing, it would have a 33-mil yaw upward. After 
burning, it would of course move in a parabola 
instead of a straight line. 

In the presence of the overturning moment, the 
up yaw caused by the gravity drop lifts the nose, 
inducing a precession first to the right and then 
down. The process is slow because the magnitude of 
the yaw causing it starts at zero and builds up 
slowly, but by the end of burning the rocket will be 


292 


EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 


somewhat to the right of the launcher line and, if 
the burning time is long enough, may drop well 
below the point where gravity alone would have 
taken it. This effect is calculated in reference 2 
and shown graphically in Figures 3 and 4, which 
give the orientation curves for the rocket axis and 
the trajectory, respectively. As in Figure 4 of Chap- 
ter 24, the quantities shown in the graphs are 
“characteristic functions” © and 4>. To obtain the 



Figure 3. Deflection of the rocket axis due to 
gravity, during burning (S — 2). 

actual angles 0 and <£ in radians for any particular 
rocket, the functions must be multiplied by the 
factor g/G for horizontal launching, or in general 
for a quadrant angle 0 O , by the factor g sin 0 O / 
(G — g sin 0 O ). C Thus, in the particular case we 
are considering, the point at the end of burning 
(f = 1.50) corresponds to 

$ gX = 1.47; $ gY = 0.6; 

= 0.56; @ ff r = 1.05. 

for the zero-length launcher. The conversion factor 
g/G — 1/30 so we calculate that the rocket is point- 

c Relations between characteristic functions and actual 
angles are given in Table 3 for all functions used in this chapter. 


ing 49 mils to the right and 20 mils below the launcher , 
and the trajectory is deflected 18.7 mils to the right 
and 35 mils downward. Here the downward de- 
flection is barely greater than it would be in the 
absence of the overturning moment, but it is appar- 
ent from the curves that with a little longer burning 
time it would become much greater. 


Table 3. Relations for converting from characteristic 
functions to actual angles.* 


Gravity: 


= ^0 
G U °' 


Mallaunching: 


dq = qt x ®q. 


Wind: 

tw - ——©IF. 
v \ 

Relations between <f> and <£ are identical. 

t x = V2 \/G 
v\ = V2G\ 

r = v d]\ 


tv = 


* All above relations assume horizontal fire. If quadrant elevation is do 
substitute G — g sin do for G and g cos do for g wherever they appear. 


Each of the curves of Figure 4 shows a minimum 
of right deflection for f « 2.8, because slightly 
before this the rocket has made one complete preces- 
sion and is ready to start heading off toward the 
right again. For higher stability factors, the rocket 
travels farther in one precession, and the gravity 
deflections for a given burning distance are some- 
what less. 

Curves giving the deflection of the center of 
mass from the range line throughout burning are 
also given in reference 2, but in most actual cases 
where the total flight distance is considerably 
greater than the burning distance, this deflection 
may be neglected, and the trajectory angle at the 
end of burning will give the final deflection with 
sufficient accuracy except for drift effects. 

After burning ceases, the curves of Figures 3 and 
4 are no longer applicable; the rocket tends to settle 
into the position where the yaw to the right pro- 
duces enough precession to cause it to follow the 
trajectory, as explained in Chapter 21. 


MOTION DURING BURNING 


293 


0 0.1 0.2 0.3 0.4 0.5 , 0.6 0.7 0.8 



Figure 4. Deviation of the trajectory due to gravity, during burning (S = 2). 



294 


EXTERIOR BALLISTICS OF SPUN-STABILIZED ROCKETS 


25.3.3 Effect of Mallaunching 

One of the most important factors in spinner 
motion, and the most difficult to control, is mal- 
launching. The term “mallaunching” is used tech- 
nically to denote any angular velocity, about a 
transverse axis, which the rocket acquires during 
launching. Such angular velocities may be pro- 
duced by gravity (tip-off), a faulty launcher, 
dynamic and static unbalance of the round, elliptical 
bourrelets, or jet malalignment. 

Because the effect of mallaunching in deviating 
the trajectory occurs almost entirely in the early 
part of burning before the velocity and the aero- 
dynamic forces become large, a fairly satisfactory 



Figure 5. Deflection of the rocket axis due to 
mallaunching, during burning (S = 2). 

treatment of it can be obtained by assuming that 
no aerodynamic forces act on the rocket. References 
3 and 4 contain this analysis. The more general 
case where the effect of the overturning moment 
is included is discussed in reference 2, and both 
cases are treated in Exterior Ballistics. 1 

If one assumes that the launcher is absolutely 
rigid and that there is no friction, malalignment, or 
unbalance, the angular velocity produced by tip-off 
is computed easily by considering the gravity torque 
acting on the rocket, supported on its rear bourrelet, 
during a time equal to that between the arrival of 
the front and rear bourrelets at the end of the 
launcher. The resulting equations are given in 
reference 4 and are identical with those for finners 
because the gyroscopic forces can produce no sig- 
nificant effect in so short a time. Practical launchers 


are not absolutely rigid, and their reaction on the 
round may impart to it either more or less angular 
velocity than the simple theory would predict. It is 
this variation in mallaunching that produces the 
sometimes rather large discrepancies in centers of 
impact among different launchers. 

If, on leaving the launcher, a rocket receives an 
angular velocity throwing the nose downward, for 
example, it responds in the manner that we have by 
now come to expect, changing the downward mo- 
tion into motion to the left. Here, however, we 
have to do, not with a precession, which is the 
response to the continued action of a force, but 
with a nutation, which is roughly 4$ times more 
rapid than a precession . The nose moves in a tight- 
ening spiral d so that virtually all the change in 
orientation occurs in the first nutation, as shown by 
Figure 5 in terms of symbols similar to those of 
Figure 3, except that we must rotate the figure 
clockwise 90 degrees in order to apply to tip-off. To 
get the actual angles, we multiply the tabulated 
functions by the factor t\ = y/2 X/G; e the result is 
expressed in angular units per unit of mallaunching 
velocity. For our hypothetical HCSR, the factor 
is 0.50 for horizontal fire. 

Using the curve for the zero-length launcher, we 
find that by the end of burning (f = 1.5) the rocket 
has completed 2 loops on its spiral and has 
coordinates 

Qqx = 0.3; 4> a y = 0.25; 

corresponding to an orientation 0.15 degree (or 
mil) below and 0.125 degree (or mil) left of the 
launcher line for an initial angular velocity of 1 
degree (or mil) per second. 

After the end of burning, in the absence of aero- 
dynamic forces, the nose would move in a circle 
having the same center and radius as the spiral had 
when the thrust ceased. With the overturning 
moment acting, this nutation will, of course, be 
superimposed on the precession. 

The direction of the trajectory during burning is 
given similarly in Figure 6. f Again we turn the 

d The reader may recognize it as a Cornu spiral, which 
gives another representation of the Fresnel integrals which 
appear so frequently in the theory of both finner and spinner 
trajectories. 

e As before, we use G — g sin do in place of G if the quadrant 
elevation is greater than zero. 

f Figure 5 represents the vacuum case, but Figure 6 includes 
the aerodynamic overturning moment, the effect of which is 
quite small in this case. 


MOTION DURING BURNING 


295 


© 


qx 


/ 

\ 



Figure 6. Deviation of the trajectory due to mallaunching, during burning (S — 2). 


figure through 90 degrees to apply to tip-off and 
examine the case £ p = 0, obtaining 

e qX = 0.188; <d qY = 0.207. 

The conversion factor is again 0.50, so that the 
trajectory angles are 0.094 mil down and 0.104 mil 
left for each mil per second of initial angular velocity. 
Thus it requires a tip-off of 180 mils per second to 
more than offset the approximately 19 mils right 
deflection which we calculated for the gravity 
effect. In practice, longer launchers are used, re- 
ducing the gravity effect relative to the tip-off effect, 
and the tip-off is large enough (approximately 100 
mils per second for the 5.0-in. GPSR g and 2 or 3 
times this for some rockets), so that it usually pre- 
dominates, and the rocket has a left orientation 
g Described in Chapter 20. 


throughout burning and drifts steadily to the left . h 

In discussing dispersion we shall be interested in 
the magnitude of the trajectory deflection without 
regard to direction for various launcher lengths. 
Measuring the radii from the origin to the f = 1.5 
points on the three curves of Figure 6, we obtain 

r, = 0; p « 0; 0 = 0.278; 6/q = 0.139; 

= 0.2; p « 5 ft; 0 = 0.178; 0/q = 0.089; 

= 0.3; p « 11 ft; © = 0.142; 6/q = 0.071; 

where 6/q is the actual trajectory angle at the end of 
burning for unit mallaunching. 

h The calculated orientation at the end of burning, analyzed 
into gravity and tip-off effects are tabulated for several 5.0-in. 
spinners in reference 5. 


296 


EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 


25 3 4 Wind Effect 

One more factor of importance during burning is 
the wind, which may alter the trajectory sig- 
nificantly. For the component of wind along the 
range line, the effect is nonlinear and quite com- 
plicated, 6 so we shall discuss only the cross-wind 


to unit wind velocity. It will be noted that a 
positive wind increases the gravity drop in all cases, 
but the lateral deflection starts downwind and then 
reverses if the burning continues long enough . After 
burning, the deflection is naturally downwind be- 
cause of the downwind component of the drag just 
as in the case of finners. 



Figure 7. Deflection of the rocket axis due to cross wind during burning (S — 2). 


effect. If a wind is blowing across the launcher 
from left to right, the effect is essentially as if the 
rocket were launched into still air with a yaw to the 
right. Hence an overturning moment exists because 
of the wind, and the rocket precesses clockwise as 
would be expected. Nutation is of little importance 
in this motion, and the deflection is slow and spread 
out through the whole of burning instead of taking 
place mostly in the first nutation as in the case of 
mallaunching. Also in contrast to the mallaunching 
effect, it is relatively insensitive to launcher length. 

The characteristic functions for cross wind are 
plotted in Figures 7 and 8. The conversion factor 
in this case is l/v x = the results applying 


Numerical values for our hypothetical example 
are (using f p = 0) 

&wx = 1.19; $1 vy = 1.63; 

©tfx = —0.03; = 0.87. 

Using the conversion factor 2.18 X 10 -3 , we find that 
at the end of burning a cross wind toward the right 
of 1 fps will tip the rocket axis 2.6 mils left and 3.6 
mils down, and deviate the trajectory 0.065 mil 
left and 1.9 mils down. The very small value of the 
lateral deviation is obviously an accidental result of 
the particular burning distance chosen. Increasing 
or decreasing the burning distance by a factor of 
2 would increase the deviation more than tenfold. 



MOTION AFTER BURNING 


®wx 


297 


- 0.5 - 0.4 - 0.3 - 0.2 - 0.1 0 0.1 0.2 0.3 



Figure 8. Deviation of the trajectory due to cross wind during burning (S = 2). 


The calculations do show, however, that wind sensi- 
tivities of 1 .5 to 2 mils per fps are obtained for low- 
stability spinners, so that gusty winds varying in 
velocity by only 5 or 10 fps can easily double the 
dispersion. As the stability factor increases, the 
characteristic curves for cross-wind effect hug the 
vertical axis more and more closely, and the vertical 
deflections also decrease, so that, if the dispersion 
produced by variable cross wind is to be kept low, a 
high stability factor is essential. 

25 4 MOTION AFTER BURNING 

In all of the foregoing discussion of motion during 
burning, the only aerodynamic effect which has 


been assumed to be acting is the overturning 
moment. This is permissible because all other aero- 
dynamic effects are smaller and do not make them- 
selves felt because of the short time involved. After 
burning, the times involved are, in general, con- 
siderably longer and virtually all the forces and 
torques may have observable consequences. 


Gravity 


As has already been mentioned, the primary 
effect of gravity is in producing curvature of the 
trajectory so that the rocket assumes an equilibrium 
yaw to the right. In addition, for high-angle fire, it 
causes large changes in velocity so that the aero- 


298 


EXTERIOR RALLISTICS OF SPIN-STABILIZED ROCKETS 


dynamic forces, and hence also the stability factor, 
vary widely in different portions of the trajectory. 

2542 Drag 

The drag force reduces the velocity gradually 
without affecting the spin. In the absence of other 
factors, it would gradually increase the stability, 
but the much larger changes in velocity caused by 
gravity make its effect relatively insignificant. 

2543 Lift and Magnus Force 

The equilibrium yaw to the right after burning 
causes the cross-wind force to be directed toward the 
right and the Magnus force to be directed up- 
wards, 1 and both forces produce drifts. The 
theoretical treatment of these effects is not very 
satisfactory, and the reader is referred to Exterior 
Ballistics 1 for quantitative details. We would ex- 
pect, however, that the drift to the right would 
be approximately proportional to the equilibrium 
yaw angle and hence [from equation (24) of Chapter 
21] proportional to the angular velocity of spin for a 
given quadrant angle. It is proportional also to the 
flight time and to the angle of elevation. 7 For the 
5.0-in./5 HCSR fired at 45-degree elevation, the 
drift amounts to approximately 34 mils. 

The Magnus force is proportional to the spin 
velocity and to the broadside area (hence, for a 
given caliber, proportional to the length L), so that 
the soaring effect, which increases the range, should 
be proportional to the factor s 2 L. This soaring effect 
is difficult to separate from other effects, but it 
appears to increase the maximum range of the 3.5-in. 
spinner by about 5 or 10 per cent. 

25.4.4 Spin Deceleration Moment 

The spin deceleration moment, in addition to its 
obvious role of reducing the spin, tends slightly to 
increase the amplitude of the nutations. This can be 
understood by noticing that its effect is directly 
opposite to that of the jet force in accelerating the 
spin during burning, so that it tends to move the 
rocket outward along the Cornu spiral of Figure 3. 
The effect is small, but not insignificant, for we 

* This is exactly true only if the Magnus moment is zero. 


shall see that damping the nutations is all-important 
in achieving stability in high-angle fire. 

25 4 5 Damping Moment 

In analogy with finners, the damping moment 
serves to remove energy from the nutations. In this 
role, however, it is overshadowed for spinners by 
the Magnus moment. 

25.4.6 Magnus Moment 

The most obvious effect of the Magnus moment 
is to alter the equilibrium yaw so that it is not 
directly to the right but is below or above this 
position according to whether the point of applica- 
tion of the Magnus moment is ahead of or behind 
the center of mass. Its most important role is in 
connection with stability, as discussed in the fol- 
lowing section. 

2 5 5 STABILITY 

The term “stability” has a rather wide variety of 
meaning. As applied to spinning rockets, it usually 
means that the yaw is small during the whole flight 
and undergoes no sudden changes. Small yaw is 
necessary in order to keep the drag low, to avoid 
losses in range and striking velocity, in order to 
minimize dispersion, and in order to have the rocket 
strike nose first as required for proper fuze operation. 
In Chapter 21 we noted that one condition necessary 
for stability is that the gyroscopic forces, expressed 
by the stability factor S, be sufficiently large. If, 
for example, S = 0.96, the nutation amplitude is 
multiplied by 3.5 every nutation, or by 525 every 
five nutations, and the nose of the rocket is very 
soon traveling in a spiral of radius comparable with 
the length of the round with a yaw that may be 30 
degrees or more. In practice the only feasible 
method of increasing stability is by increasing the 
spin. If the stability factor is high enough to get the 
rocket safely to the end of burning, no later trouble 
from this source will develop, since the drag reduces 
the velocity faster than the spin deceleration 
moment reduces the spin. 

When a spinner is fired at an elevation angle too 
high for its rate of spin, an entirely different type of 


STABILITY 


299 


instability sets in at or somewhat beyond the peak 
of the trajectory. The yaw builds up suddenly to a 
very large value, the rocket emits a noise which has 
come to be known among range workers as a 
“wow-wow,” and the projectile strikes the ground 
approximately broadside and usually considerably 
to the left of its normal impact point. This behavior 
occurs because the gyroscopic stability prevents the 
rocket from aligning itself promptly with the 
rapidly changing direction of the trajectory, so 
that the yaw exceeds a certain critical value. What 
determines the critical yaw we shall see presently. 

We have seen in Chapter 21 that a spinner is able 
to follow its curved trajectory because it has an 
equilibrium yaw to the right so that the overturning 
moment makes the nose precess downward. As 
indicated by equation (24) of Chapter 21, the mag- 
nitude of this equilibrium yaw for any point on the 
trajectory is proportional to the component of 
gravity normal to the trajectory and inversely pro- 
portional to the velocity. Both these factors vary 
in such a way as to make the equilibrium yaw a 
maximum at the peak of the trajectory and critically 
dependent on the quadrant elevation. As an 
example, 7a a rocket which has an equilibrium yaw of 
1 degree at the end of burning for horizontal fire 
may have the following values at the summits of 
high-angle trajectories: 

Degrees 

Angle of elevation 30 40 50 55 60 

Equilibrium angle of yaw 2.30 3.6 6.4 9.4 14.5 

These values are probably a fairly good approxima- 
tion to the equilibrium yaws of the 3.5-in. spinner, 
but are too high for most of the 5.0-in. barrage 
spinners. 

Our assumption that the overturning moment is 
proportional to the yaw or to the sine of the yaw is 
obviously false for large yaws. Long before the yaw 
becomes 90 degrees, this moment goes through a 
maximum and then usually decreases to zero and 
changes into a righting moment. A spinning rocket 
for which the overturning moment is negative will 
apparently have its equilibrium yaw to the left, 
and hence the earliest explanation of the “wow- 
wows” was as follows.* 

As the projectile approaches the summit, the 
tangent to the trajectory turns more and more 
rapidly, and the projectile must yaw farther and 
farther right so that the aerodynamic moment will 

j This explanation is derived in greater detail in reference 7. 


be large enough to cause tjhe nose to precess down- 
ward at the same rate that the trajectory turns. 
Eventually it reaches the angle corresponding to 
maximum overturning moment, and for greater 
yaws the moment decreases; then it is impossible 
for the axis of the projectile to turn as rapidly as the 
trajectory does. The nose continues to precess down 
and to the right, but the trajectory turns downward 
much more rapidly so that the yaw increases to the 
point where the aerodynamic moment reverses sign 






















„ 
















5>5 HCSfl 

! 













5710 

CnS 

R - 































































/ 

















—L 




\ 

t 


/ 








• ^ 



! 




\ 

X 





'' 







k 4 





\ 


0 

10 

20 

3 

O " 


O 

50 

1 60 

7 

) 

0 

8 

O 

90 









' 


L 







































... 


\ 


















\ 

\ 

1 



















\ 


















\ 


Figure 9. Variation of overturning moment co- 
efficient with yaw for typical 5.0-in. spinners. C M is 
proportional to the overturning moment divided 
by the square of the velocity. 

and becomes a righting moment. As long as this 
moment is no larger than the maximum overturning 
moment, the nose of the rocket precesses upward 
and back to the left at a relatively slow rate; but if, 
as the velocity and yaw increase, the righting 
moment becomes large enough, there is a new 
equilibrium yaw position in which the rocket has a 
large left yaw. Its axis then spirals around this new 
equilibrium position with an amplitude that is very 
large because the initial position was so far from the 
equilibrium position. 

This theory explained the qualitative behavior 
very well, but it broke down completely as soon as 
wind and water tunnel data and especially yaw 
camera data began to become available. Thus it 
was found that most spinners become unstable at an 
equilibrium yaw in the neighborhood of 10 degrees, 
whereas the yaw for which the overturning moment 
is a maximum is always considerably greater than 
this value. The actual variation of overturning 




300 


EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 


moment with yaw differs greatly for different rock- 
ets, as can be seen in Figure 9, where the quantity 
plotted is the overturning moment coefficients 
The true nature of the instability was first re- 
vealed by yaw camera records * 1 * such as those in 
Figure 10. This record shows the variation over an 
interval of about 10 seconds in the angle between 
the axis of the rocket and the rays of the sun. The 
oscillations whose amplitude is increasing nearly 
exponentially are the nutations. The time scale is 
defined by the 0.14-second period of the nutations. 


listics, 1 that the Magnus torque is the only aero- 
dynamic force which, averaged over a nutation, 
can add or subtract a significant amount of energy, 
and that it is responsible for the instability. 

The direction of the Magnus force is perpendicular 
to the trajectory and to the plane of yaw, and its 
point of application depends rather critically on the 
yaw. For very large yaws, it is probably at the 
center of figure of the rocket, which is usually 
slightly back of the center of mass, but for small 
yaws it is usually ahead of the center of mass. 



i 

1 

Figure 10. Yaw camera record for spinner which becomes unstable because of negative damping of the 
nutations at the peak of the trajectory. 


Evidently the instability is due to a building up of 
the nutations rather than any change in the preces- 
sional motion . Records covering the early part of an 
unstable trajectory show that for several seconds 
after launching the nutations are damped in the 
same way that they are throughout all of a normal 
trajectory. Apparently, when the yaw exceeds a 
certain critical value, something begins to pour 
energy into the nutational motion . It was shown in 
reference 11, and in greater detail in Exterior Bal- 

k Curves are reproduced from a local memorandum 8 based 
on data from a National Bureau of Standards report 9 on wind 
tunnel measurements and on various reports on high-speed 
water tunnel measurements by the CIT Hydraulic Machinery 
Laboratory. 

1 The yaw camera is described in Field Testing of Rockets, 10 

one of the CIT OEMsr-418 final reports. 


The Magnus force probably varies fairly closely with 
the sine of the angle of yaw, but, because of the 
shift in its point of application from ahead of to 
behind the center of mass, the Magnus moment is 
positive (i.e., overturning) for small yaws and nega- 
tive for large yaws, and its maximum positive value 
may occur for yaws of only a few degrees. The 
damping effect of the Magnus moment is easily 
understood from Figure 11. In part A of the 
figure is plotted a hypothetical variation of Magnus 
moment with yaw, and in B we consider the magni- 
tude and direction of the Magnus torque during a 
single nutation for two cases where the equilibrium 
yaws correspond to points A, B, and C. 

To simplify the figure, the precessional motion is 
omitted, and as usual the rocket’s varying orienta- 


DISPERSION OF SPINNERS 


301 


tion is represented by the curve (a circle) traced by 
its nose as we look along the trajectory in the 
direction of motion, the difference between this and 
previous figures being that the Z axis is now along 
the trajectory instead of along the launcher. The 
straight arrows represent the torque at various 
points during the nutation, their magnitudes being 
obtained from the upper curve and their directions m 
being at right angles to the line representing the yaw 
(i.e., the line from T to the point on the circle). It 
is evident that for small yaws (points A and C) the 
net effect of the torques is to oppose the motion and 
hence damp out the nutations, whereas for large 
yaws (point B), the net effect is in the same direc- 
tion as the motion, and the damping is negative. 



Figure 11. Diagram illustrating the effect of the 
Magnus moment during one nutation. 


The derivation of the exact critical yaw which 
separates positive from negative damping is some- 
what involved, but the result is shown in part A 
of the figure. One might expect it to be at the exact 
peak C, but it is displaced slightly beyond by the 
fact that the vectors representing the torques are 
not all parallel. 

When, near the peak of the trajectory, the 
equilibrium yaw exceeds the critical yaw, the nuta- 
tion amplitude begins to build up slowly so that it 
will be somewhat beyond the peak that the actual 
“wow- wows” begin. In fact, if the angle of ele- 
vation is only very slightly too large, the equilibrium 
yaw may decrease below the critical point on the 
descending part of the trajectory before the nuta- 

m These arrows are not conventional torque vectors but 
point in the direction which the torque tends to move the nose 
of the rocket. 


tions have built up signifipantly, so that nothing 
noticeable happens even though the damping was 
negative for a time. 

25 5 1 Effect of Wind on Stability 

From the preceding analysis we can immediately 
derive one important effect of down-range winds, 
the treatment of which we have omitted because of 
its complexity. Obviously a wind in the direction 
of the motion will reduce the aerodynamic forces, 
since they depend on the relative velocity between 
rocket and air; thus a larger yaw angle will be re- 
quired to give enough precession to turn the rocket 
over the top of its trajectory, and it will become 
unstable at somewhat lower elevation angles. An 
up-range wind, on the other hand, increases the 
maximum angle of elevation at which stability over 
the trajectory peak can be retained. 

25 6 DISPERSION OF SPINNERS 

The two principal advantages of spinners over 
finners are their more convenient shape and their 
usually smaller dispersion. Their greater accuracy 
stems from the fact that the spin changes the direc- 
tion of the malalignment torque so rapidly that it 
averages approximately to zero, thus by-passing 
the barrier of gas malalignment which limits the 
accuracy of finners. The introduction of spin, how- 
ever, creates many more new problems than it 
solves, and considerable effort is required if the dis- 
persion of a spinner is to be much less than half 
that of a typical well-designed finner. 

Dispersion of spinners may arise from any of the 
following causes: 

1. Variation in wind velocity. 

2. Variation in tip-off. 

3. Out-of-roundness of the bourrelets. 

4. Static and/or dynamic unbalance. 

5. Malalignment. 

We have already treated the wind effect and have 
seen that it can be reduced by using longer launchers 
or by increasing the stability factor. Causes 2 and 3 
lead to dispersions which are smaller than those 
caused by 4 and 5, and which depend on such 
variables as launcher length and burning time in 
the same way as the latter, so we shall not discuss 
them here. If the rounds have been carefully bal- 


302 


EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 


anced, the out-of-roundness might become im- 
portant, however; it is discussed further in Exterior 
Ballistics. 1 

256 1 Unbalance 


A symmetrical body is said to be statically un- 
balanced when its center of mass does not lie on its 
axis (see Figure 12). For spinners, we will take the 



MOTOR TUBE WITH PURE DYNAMIC BALANCE AND 
PURE STATIC BALANCE (DYNAMICALLY BALANCED) 



MOTOR TUBE WITH PURE DYNAMIC UNBALANCE AND 
PURE STATIC BALANCE (DYNAMICALLY UNBALANCED) 



MOTOR TUBE WITH PURE DYNAMIC BALANCE AND 
PURE STATIC UNBALANCE (DYNAMICALLY UNBALANCED) 


Figure 12. Types of unbalance of a spinner. The 
small weights W represent overweight sections of 
the tube such as occur from inequality in wall 
thickness. The center of gravity is the point CG. 

axis to be that of the bourrelets since this is the one 
about which the rocket is forced to rotate in a rigid, 
snug-fitting launcher. As a quantitative measure 
of static unbalance we will take the angle u s defined 
by the ratio of the distance between the center of 
mass and the bourrelet axis to the distance between 
the bourrelets. 

In a spinner with this sort of unbalance," the 
sides of both bourrelets toward which the center of 
mass is displaced will exert a (centrifugal) force on 

n Assuming the usual case in which the center of mass is 
between the bourrelets. 


the launcher guides, which thus must exert a 
(centripetal) reaction force to maintain the rotation 
about the bourrelet axis. After the front bourrelet 
clears the launcher, it is no longer subject to this 
reaction force, but the rear bourrelet is. The result 
is a transverse angular velocity in the direction of 
the unbalance. (Actually the direction of the unbal- 
ance is changing constantly, so that the direction of 
the angular velocity is a sort of average of the 
directions which the unbalance had when the two 
bourrelets cleared the launcher.) 

Dynamic unbalance ° is a slightly more com- 
plicated concept and is entirely independent of 
static unbalance, that is, of the position of the 
center of mass. In the absence of external forces, 
the only stable rotational state of a rigid body is 
rotation about an axis which makes its moment of 
inertia either a maximum or a minimum. There are 
in general three such axes at right angles to each 
other, and they are known as the principal axes of 
inertia. A perfect spinner would have its principal 
axis corresponding to minimum moment of inertia 
coincident with the bourrelet axis, but in general 
there is a small angle ud between them, which we 
shall take as the measure of the dynamic unbalance. 
If the launcher is tight enough and rigid enough to 
constrain the round to rotate about its bourrelet 
axis, the dynamic unbalance creates centrifugal 
forces causing opposite sides of the two bourrelets 
to press against the launcher, and, when the 
rocket is freed, the transition from rotation about 
the bourrelet axis to rotation about the axis of 
inertia produces a transverse angular velocity, i.e., 
a mallaunching. 

The calculation of the amount of mallaunching 
with a real launcher is extremely difficult. However, 
since experiment has shown little difference between 
dispersions produced by light flexible launchers and 
heavy rigid ones, it is probably sufficient to make 
calculations for an absolutely rigid launcher. This 
is much simpler and is usually what is done. A very 
elementary calculation will give us an estimate of 
the mallaunching in this case. We need merely 
consider the vector s p representing the spin angular 
velocity at the moment of launching (in this ap- 
proximation we assume that the constraint is 
removed from both bourrelets simultaneously) and 
resolve it into two components, one parallel to and 

° Ordinarily, in the literature, a spinner is said to be dy- 
namically unbalanced when it has either or both of the two 
types of unbalance. 


DISPERSION OF SPINNERS 


303 


one perpendicular to the axis of inertia. The latter 
is the transverse angular velocity. 

q = s p sin Ud ~ s p Ud- 

This derivation neglects the details of the interac- 
tion between the launcher and the two bourrelets 
and cannot be expected to give an exact answer, but 
it does show correctly that the mallaunching due to 
unbalance is proportional to the spin rate at launch- 
ing. When the effect of static unbalance is con- 
sidered also, 12 we find that the mallaunching is given 
to a good approximation by 

q = is p u, 

where 

u = Vu D 2 + 2 Us 2 . 

The method of combining dynamic and static un- 
balance by the square root of the sum of the squares 
assumes that they are randomly oriented relative to 
each other, and the factors 2 and % come out of 
more complicated analysis. 

We are now in a position to apply the mallaunch- 
ing formulas to the determination of the dispersion 
to be expected as a result of dynamic and static 
unbalance. At the end of Section 25.3.3 we cal- 
culated the deflection of our hypothetical HCSR 
at the end of burning for unit mallaunching from 
launchers of three different lengths. The mal- 
launching due to unbalance for a truly zero-length 
launcher would be zero according to our calcula- 
tions, but we can apply the zero-length solution 
approximately to a very short launcher, say 1 ft 
long, which is just the bourrelet spacing for the 
5.0-in./5 HCSR. 

For the spin at launching, we have 
Sp = yV2Gp = 46 Vp 

= 46 radians per second for p = 1 
= 103 radians per second for p = 5 
= 152 radians per second for p = 11. 

If we assume that the total unbalance u is 0.001 
radian, the angular velocities of mallaunching 
produced by the three launcher lengths are re- 
spectively 0.0345, 0.077, and 0.114 radians per 


second. Hence, from the values of 6/q at the* end of 
Section 25.3.3, we calculate the deflections to be 

4.8 tails for p = 1 ft; 

6.9 mils for p = 5 ft; 

8.1 mils for p = 11 ft. 

From these three values it appears that dispersion 
due to unbalance is small for very short launchers, 
and that for launchers of practicable length it is a 
very slowly increasing function of launcher length. 
These same conclusions are arrived at in a different 
way in Rocket Design , 12 and they appear to be sup- 
ported by the experimental data. Since very short 
launchers are not practicable, as we shall see, one 
more conclusion can be drawn: namely, that if the 
dispersion of the HCSR is to be much less than 10 
mils, considerable care must be taken in balancing 
it. Thus consider the 6.9-mil figure. To get the 
lateral dispersion for firing at 45-degree elevation, 
this is multiplied by 2/n (because of the random 
orientation of the unbalance) and by \/2 (to cor- 
rect for elevation angle), and the result is 6.2 mils. 
The actual dispersion of the HCSR is about 50 per 
cent greater than this, so an unbalance of only 
0.001 radian (i.e., 1 mil) will account for most of it. 
The actual magnitudes of unbalance exhibited 
by production line spinners are given in Table 4. 


Table 4. Effect of dynamic unbalance on dispersion of 
spinners. 


Type of rocket 

3. 5-in. /4 
GPSR 

5.0-in./10 

GPSR 

5.0-in. /14 
GASR 

Number of rockets 

Purely static unbalance (dis- 
placement of center of 
gravity from geometric 

100 

93 

129 

axis in in.) 

Purely dynamic unbalance 
(angular deviation of dy- 
namic axis from geometric 

0.0035 

0.0137 

0.0094 

axis in mils) 

0.53 

0.88 

0.97 

Mean deflection (mils) 

Mean lateral or vertical dis- 

5.9 

7.8 

6.2 

persions (mils) 

4.2 

5.5 

4.4 


25.6.2 Effect of Jet Malalignment 

The simplest treatment of jet malalignment for 
spinners is to consider its effect as being equivalent 
to a mallaunching. This is possible because only for 
a very short time after it becomes free of the launcher 


304 


EXTERIOR RALLISTICS OF SPIN-STARILIZED ROCKETS 


is the rocket spinning slowly enough for the effect 
of malalignment to be significant. The relationship 
between malalignment and equivalent transverse 
angular velocity of mallaunching, derived in Ex- 
terior Ballistics , lb is approximately 

_ VWh 
q ~ \KW 2 ’ 

where R m is the jet malalignment in feet and E is a 
dimensionless function of the launcher length and 
certain constants of the round, which has been cal- 
culated and tabulated in Exterior Ballistics. 1 For 
our purpose, we may take E to be given with suffi- 
cient accuracy by p 

1 

E = VI + 20 p/fyxfc 2 ' 

Thus E is equal to unity for a zero-length launcher 
and decreases rapidly with increasing launcher 
length, corresponding to the fact that a given mal- 
alignment results in less mallaunching on longer 
launchers, as we would expect, since the rate of 
spin on leaving the launcher is higher. 

For example, in an HCSR of the characteristics 
given in Table 1, with a malalignment R m of 0.001 
in. (0.0000833 ft), the deflections of the trajectory 
from launchers of length 0 and 5 ft, will be as fol- 
lows, in terms of equivalent mallaunching effect: 


1. V 

ft 

0 

5 

2. R m 

ft 

0.0000833 

0.0000833 

rJfV 2 

ft" 1 

64 

64 

4. E 


1.00 

0.238 

5. q 

radians per second 

5.33 

1.29 

6. 0/q 

seconds 

0.139 

0.089 

7. 9 

mils 

0.74 

0.115 


In this tabulation, the first five lines are simply 
data for and evaluation of the preceding equations, 
to give, in line 5, the transverse angular velocities 
of mallaunching equivalent, in deflection effect on 
the trajectory, to the jet malalignment. These 
quantities are multiplied by the data in line 6, 
which are the trajectory deflections (angles) per unit 
angular velocity of mallaunching listed at the end of 
Section 25.3.3. The figures in line 7 give the 
trajectory deflection angles, in mils per thousandth 
of an inch malalignment. 

p E differs by a constant factor from the function | E ^ [ of 
Exterior Ballistic s. 1 


This calculation shows us that for very short 
launchers, malalignment for spinners would be as 
serious as for some fin-stabilized rockets (the HVAR 
deflection, for example, is 1.0 mils per 0.001-in. 
malalignment). Hence spinner launchers are made 
just long enough so that the dispersion due to mal- 
alignment is considerably smaller than that due to 
unbalance. Further increasing the launcher length 
gives little improvement because of the very slow 
change in unbalance effect with launcher length. 

25 6 3 Optimum Spin 

It has been mentioned before, but probably 
cannot be too strongly emphasized, that the attain- 
ment of high accuracy, say 5 mils or better, prob- 
ably requires the use of higher rates of spin than 
those used in any present ground-fired spinners. 
Higher spin would reduce the high sensitivity of the 
present rounds to cross winds. Analysis shows that 
the wind sensitivity (i.e., the trajectory deviation 
per unit cross wind) is approximately inversely 
proportional to the stability factor and hence, other 
factors being the same, inversely proportional to the 
square of the spin velocity. When the spin is in- 
creased, dispersion due to malalignment decreases. 
Dispersion due to mallaunching of constant mag- 
nitude also decreases, but, as we have seen, the 
magnitude of the mallaunching is likely to increase . 
This can probably be mostly compensated by the 
reduction in launcher length which reduced mal- 
alignment effect makes possible. With the wind 
effect reduced, it would then be profitable to take 
greater care in balancing the rounds. 

On the other hand, such high-stability rounds 
will not follow a rapidly turning trajectory, so that, 
in cases where it is required that they do so, one 
must probably be content with present dispersions. 

25 7 SPINNER RANGE CALCULATIONS 

For a perfectly launched spinner, the range 
calculations would differ little from those for a 
finner. In each case the starting point is the theory 
of the equivalent shell. The estimation of air drag 
is easier for a spinner because of the usually simpler 
shape and lack of lugs and fins, but the drag is a 
function of yaw angle and, since the spinner does 
not have an average zero yaw, this relationship 


O 


TRAJECTORIES OF SPINNERS FIRED FORWARD FROM AIRPLANES 


305 


must be known . The difficulty with purely theoret- 
ical range calculations is that spinners are almost 
never perfectly launched, and the direction and 
magnitude of the mallaunching depend upon the 
propellant temperature and upon the particular 
launcher. Thus the only feasible way to construct 
range tables is to start with experimental data 
under standard conditions and use the theory to 
make corrections to other conditions. This is the 
system adopted in reference 13, and for details of 
the procedure, the reader is referred to this report. 
The basic theory is given in Exterior Ballistics. 1 
Obviously the same remarks apply to the mean 
deflections as to the ranges. 

258 TRAJECTORIES OF SPINNERS 
FIRED FORWARD FROM AIRPLANES 

All the characteristic functions for spinner motion 
which are tabulated in this chapter were calculated 
on two assumptions: (1) constant stability factor 
and (2) proportionality of overturning moment to 
yaw angle. The reason is that the solutions of the 
equations of motion for the more general case are 
not possible in terms of functions with which 
mathematicians are familiar and can be evaluated 
only by numerical methods. 

In ground firing, this rather restricted theoretical 
treatment covers many cases of interest. Thus, if 
we stay well below the velocity of sound, the 
stability factor is very nearly constant during 
burning and changes very slowly thereafter. Yaws 
do not exceed about 10 degrees, in which range the 
nonlinearity of the overturning moment is not great 
enough to alter the motion significantly. In no 
case, however, are these conditions true in air- 
craft firing, so that a rather comprehensive pro- 
gram of computations with a differential analyzer 
may be required before sighting tables such as those 
for fin-stabilized aircraft rockets can be made. 

When spinners are fired forward from airplanes, 
they are subjected to large aerodynamic forces as 
soon as they clear the launcher, while their spin is 
still small. As a result, their stability factor is below 
the critical value, and the yaw and transverse 
angular velocity tend to increase rapidly. Usually 
the spin increases to the point where the rocket be- 
comes stable again so that the yaw is damped out, 
but two factors may prevent this. (1) The over- 


turning moment coefficient increases considerably 
at approximately the velocity of sound, thus re- 
ducing the stability factor so that the rocket may 
not be stable even at the end of burning when the 
spin is greatest. Thus a spinner having S = 6 for 
ground firing may drop to S = 2.5 at the end of 
burning in forward firing at high airplane speeds. 
(2) The yaw may build up to the point where the 
nutations become negatively damped by the Mag- 
nus moment. Even though the rocket may recover, 
it is likely to acquire a rather large deflection during 
its period of instability, so it is desirable to reduce 
the duration of this period as much as possible. 

From the expression for the stability factor, equa- 
tion (23) of Chapter 21, it is seen that the most 
convenient ways to increase the stability of a rocket 
are to increase its spin or to decrease its transverse 
radius of gyration, that is, its length. Because the 
spin is limited by the centrifugal force which the 
grain can stand, it was found necessary in adapting 
the 5.0-in. spinner to forward firing to change both 
the spin and the length . It may also be possible to 
increase the stability by moving the center of mass 
of the rocket forward, thus reducing the overturning 
moment coefficient /z. It must not, however, be 
moved so far that the Magnus moment reverses, 
or the rocket will be unstable after burning because 
of the negative damping of the nutations. 

The requirements for a spinner to be fired side- 
ways from an airplane are similar to those for 
forward firing. If the rocket maintains its orienta- 
tion during burning, the speed will soon build up to 
the point where yaw with respect to the air is small 
even though it may have been nearly 90 degrees at 
the start. The rocket will then be stable. The 
condition that its orientation be unchanged is that 
the transverse angular velocity build up slowly and 
be small when the rocket has become stable. This 
requires a large spin velocity and a low overturning 
moment at large yaw. 

When fired backward from aircraft, the rocket is 
for a time moving base first through the air with 
decreasing speed so that it becomes stable con- 
siderably sooner than when fired forward. No 
ballistic calculations are possible during this period 
because the normal airflow from base to nose along 
the rocket is completely disrupted by the jet. It is 
difficult to say what is required in this case, since no 
experimental data are available, but it appears that 
high spin will be desirable here also. 


306 


EXTERIOR BALLISTICS OF SPIN-STABILIZED ROCKETS 


259 TERMINAL BALLISTICS OF 
SPIN-STABILIZED ROCKETS ** 

There is a current impression that the underwater 
and underground trajectories of spin-stabilized rock- 
ets and shells are short because of the spin. How- 
ever a close examination of the question shows that 
the presence or absence of spin should be of rela- 
tively little importance in determining this feature 
of the terminal ballistics of the projectile. All im- 
portant differences between the behavior of fin- 
stabilized rockets and spin-stabilized projectiles are 
due to other factors, such as the nose shape, the 
ratio of length to diameter, and the ratio of mass 
to cross-sectional area. It is probable that the 
terminal ballistics of spin-stabilized rockets could be 
improved, if desired, by the use of the proper nose 
shapes. 

The only ways in which spin could modify the 
terminal ballistics of a projectile are if the nutation 
and precession associated with the spin modify the 
motion, or if the spin causes the medium to exert 
additional forces on the projectile. Now gyroscopic 
effects are not evident in the usual projectile until 
it has made three or four revolutions. Over shorter 
periods it responds to applied forces in essentially 
the same way as does an unrotated projectile. 
Hence it seems clear that gyroscopic effects are 
unimportant during entry. When the projectile is 
traveling in the bubble under water or in earth, 
there will be large forces exerted at the nose and tail, 
but the total torque acting about a transverse axis 
is extremely small, as is shown by the fact that the 
angular acceleration about such an axis is small for a 
fin-stabilized rocket. Hence we should expect no 
serious gyroscopic effects on the trajectory of a 
spin-stabilized projectile, provided it has a satis- 
factory underwater or underground trajectory when 
not spinning. There might be some tendency for 
the simple circular trajectory of a nonspinning pro- 
jectile to be warped into a section of a helix, but the 
pitch would be long and the total distance traveled 
the same. It seems safe to assume that the only 
additional forces and torques of appreciable mag- 
nitude exerted by the medium, because of the pres- 

q This section is adapted from an informal memorandum by 
Leverett Davis, Jr. 


ence of spin, are a torque tending to decrease the 
spin about the longitudinal axis. 

It follows from these considerations that the 
theory of underwater ballistics discussed in Chap- 
ters 15 and 24 can be applied to spin-stabilized 
projectiles, as well as to nonspinning projectiles. 

Probably the most important factors to be con- 
sidered in getting satisfactory terminal ballistics 
in water and earth are the use of such a nose shape 
and such a ratio of length to diameter that the 
cross forces and the curvature of the trajectory are 
small. The next most important factors are the use 
of a nose shape having a small drag coefficient and 
the use of a large ratio of mass to cross-sectional 
area in order to get a small deceleration for a given 
drag and hence to get a long underwater or under- 
ground travel. 

Well-designed fin-stabilized rockets tend to be 
longer and heavier than well-designed spin-stabilized 
rockets of the same diameter, and spin-stabilized 
rockets tend to be longer than shells. These charac- 
teristics are the result of efforts to secure efficient 
rocket propulsion, low dispersion, and satisfactory 
flight. It follows that fin-stabilized rockets will 
almost always have somewhat longer and straighter 
underwater and underground trajectories than will 
spin-stabilized rounds. However the usual spin- 
stabilized rocket has such a nose shape that its 
underwater trajectory is much poorer than the 
optimum set by its length and mass. In the case of 
the 5.0-in./10 GPSR, the length to diameter ratio 
is 7, and it is probable that by the use of a suitable 
nose shape the underwater behavior could be con- 
siderably improved. It may prove to be more diffi- 
cult to get a satisfactory underwater trajectory for 
the aircraft spinners, since their length to diameter 
ratio is only 5. The improvement possible will have 
to be determined by experiment. Shells are usually 
from 4 to 6 calibers long, and hence it may be diffi- 
cult to achieve a satisfactory underwater trajectory, 
particularly since the nose shape is usually chosen 
to give low drag in air and this tends to give very 
large cross forces in water. It should be noted, 
however, that the British seem to have had con- 
siderable success in designing shells having a 
relatively long underwater trajectory. 


GENERAL BIBLIOGRAPHY OF TECHNICAL REPORTS ISSUED UNDER 


DIVISION 3, NDRC 


Of the many types of technical reports resulting 
from the work of Division 3 during World War II, 
the following lists include nearly all those believed 
to be of continuing general value. As far as possible, 
the reports are listed under the contracts covering 
the activities reported. Under “Noncontract Re- 
ports’ ’ are listed those which cannot properly be 
credited or limited to any single contract. The scope 
of the work under each contract is indicated briefly 
under “Contract Numbers” and at more length in 
the Introduction to this one-volume Summary 
Technical Report of Division 3. 

In Division 3 (and its predecessor NDRC units, 
Sections C and H of Division A) the forms of re- 
ports and the machinery for their preparation, 
identification, and distribution varied with time 
and from one contract to another. Hence, there is 
some lack of uniformity in the following lists. 

Under each contract the OSRD monographs (if 
any) are listed first, followed by the final reports, 
then the interim reports, then the periodical reports 
(if any), and finally miscellaneous other types. The 
monographs are unclassified publications to be pub- 
lished by the McGraw-Hill Book Company; they 
cover theory and principles developed during the 
contract work in fields indicated by their titles. The 
final reports review and summarize the contract 
activities and results. On two contracts ultrafinal 
“summary reports” recapitulate the information 
presented in many final reports on separate pro- 
grams . 

Most of the interim reports present the status 
and results of individual research and development 
projects or programs as of the times when the re- 
ports were prepared. These were written as results 
became available, frequently in response to ex- 
pressed needs for the information. 

Under the larger contracts, periodical reports 
(weekly, biweekly, or monthly) were prepared, pri- 
marily for intracontract and intradivision informa- 
tion and control. These outlined activities and pre- 
sented test results, usually in substantial detail. 

The nomenclature used in titles was that in com- 
mon use when the reports were written; in some 
cases it has changed since then. 


IDENTIFICATION NUMBERING 

All Division 3 reports distributed through or for 
the Executive Secretary of OSRD bear OSRD 
numbers included in the lists below. In addition 
to the OSRD numbers, many of the reports also 
bear NDRC “A” numbers; all these were edited, 
reproduced, and distributed by the NDRC (Di- 
vision A) Technical Reports Section. Most of the 
reports prepared under Contracts OEMsr-273 and 
OEMsr-418 had contractors’ numbers assigned to 
them; in some cases these are the only identifica- 
tions. The “A” series reports are listed with dates 
of issue, sometimes months after preparation of the 
manuscript. In most other cases the dates listed are 
closer to the date of manuscript completion. 

AVAILABILITY 

Most of the more formal types of Division 3 re- 
ports were rather widely distributed to Army and 
Navy offices concerned with the development of 
rockets and underwater ordnance. Copies of the 
monographs will be distributed also to the Army 
and Navy. Additional copies will be obtainable only 
by purchase; the publisher has agreed to a 3334$ 
per cent discount on purchases by government 
agencies. With the few exceptions noted in the lists 
below, several hundred copies of each final report 
were distributed to the Army and Navy. Fifty to 
one hundred or more copies of all “A” series re- 
ports were so distributed, as were larger numbers 
of all except the first few of the OEMsr-418 J, 
LMC, and PMC series. Except where otherwise 
noted in the lists below, other types of reports were 
given only limited distribution. 

In connection with termination of contracts, all 
OSRD contractors were required to dispose of all 
copies of their contract reports, except for record 
copies. In the case of Contract OEMsr-418, at 
least one complete set, several nearly complete sets, 
and large numbers of surplus copies of reports were 
transferred to the custody of the Naval Ordnance 
Test Station, Inyokern, California. In the case of 
Contract OEMsr-273, collections of reports were 
transferred to the Navy Bureau of Ordnance, for 


307 


308 


BIBLIOGRAPHY 


use at the Allegany Ballistics Laboratory (now 
operated by the Hercules Powder Company) and 
to the Johns Hopkins University Applied Physics 
Laboratory. The Bureau of Ordnance took over also 
the collections of reports associated with the 
OEMsr-716 work at the University of Minnesota. 
Other contractors delivered their surplus reports 
to the Library of Congress, either directly or 
through OSRD. 

Record copies of the more formal Division 3 re- 
ports are in the files of Division 3, of the OSRD 
Executive Secretary, and of the OSRD Liaison Of- 
fice. Rather complete collections of Division 3 re- 
ports are maintained by Office of Naval Research, 
Navy Department; Research and Development Di- 
vision, War Department General Staff; Bureau of 
Ordnance, Navy Department; and Office of the 
Chief of Ordnance, War Department. 

Surplus copies were transferred to the Library of 
Congress. OSRD reports there are available to 
meet Army and Navy requests. Photostatic and 
microfilm copies of unclassified reports are avail- 
able to the public from the Office of Technical 
Services, U.S. Department of Commerce. 

Of the reports listed below, positive and negative 
microfilms of those with microfilm numbers are to 
be deposited with the Coordinator of War Depart- 


ment Libraries, as a supplement to this Summary 
Technical Report, to meet government needs. Micro- 
filmed reports are listed in the volume bibliography 
starting on page 347. 

SECURITY CLASSIFICATION 

Nearly all Division 3 reports were classified when 
issued. Most of them were confidential. As the re- 
sult of reclassification review after World War II 
ended, few of them retain this classification now. 
Some reports indicating uses and performance of 
weapons and components involved in continuing 
Service development programs remain restricted. 
Most of the reports on standardized or obsolete 
weapons and on general theory, principles, phe- 
nomena, and instrumentation have been declassi- 
fied; others may be in the future. 

RELATED REPORTS 

As indicated in the Introduction to this Summary 
Technical Report volume, work related to that of 
Division 3 was done by other divisions, the Services, 
their contractors, and other agencies (including 
some of the United Kingdom). No adequate at- 
tempt can be made in this volume to list or refer 
to the reports on this work. 


NONCONTRACT REPORTS OF DIVISION 3 


NDRC No. OSRD No. 


A-4 

8 

A-21 

24 

A-22 

316 

A-24 

320 

A-25 

323 

A-27 

311 

A-28 

345 

A-36 

466 

A-69 

691 

A-70 

673 

A-78 

769 

A-102 

943 

A-117 

1070 

A-127 

1136 

A-150 

1264 

A-166 

1359 

A-178 

1473 


Jet Acceleration of Armor-Piercing Bombs, as of March 1, 1941, C. N. Hickman, June 1941. 

The Use of Copper Balls for Measuring Pressures in Combustion Chambers, C. N. Hickman, Nov. 
1941. 

Internal Ballistics of Power Driven Rockets, E. Lakatos, Dec. 1941. 

The Mechanical Efficiencies of Rockets in Empty Field-Free Space, J. W. M. DuMond, Jan. 1942. 
Microtome Sections of Ballistite Powder, A. J. Dempster, Dec. 1941. 

Rocket Targets as of November 1, 1941, A. J. Dempster, Dec. 1941. 

The Trajectories of Target Rockets, A. J. Dempster, Jan. 1942. 

A Vacuum Press for the Extrusion of Solventless Double-Base Ballistite for Use in Rockets, J. W. M. 
DuMond, Mar. 1942. 

Jet Acceleration Tests of the 14-in. Armor-Piercing Bomb, C. N. Hickman, July 1942. 

A 4)/2-in. High-Explosive Rocket Shell for Projection from Airplanes, C. N. Hickman and L. A. 
Skinner, July 1942. 

Inspection and Testing of a %-in. Stick Powder, J. Beek, Jr., Aug. 1942. 

The Rate of Burning of Double-Base Powders and the Possible Effects of Change in Nitroglycerin and 
Total-Volatiles Content on the Burning of Jet Propulsion Tube Power, R. E. Gibson, Oct. 1942. 
Microscopic Structure and the Development of Flaws in Extruded Grains of NT Smokeless Powder, 
C. P. Saylor, Nov. 1942. 

Microscopic Examination of Extruded Smokeless Powders, C. P. Saylor, Dec. 1942. 

The Design of Granulation for Rocket Powder, J. Beek, Jr., Feb. 1943. 

Some Problems of Heat Transfer in Rockets, J. Beek, Jr., Apr. 1943. 

Studies of the 4-2-in. Chemical M ortar, A. R. T. Denues, Apr. 1943. 


BIBLIOGRAPHY 


309 


NDRC No. 

OSRD No. 

A-185 

1466 

. A-197 

1590 

A-198 

1649 

A-202 

A-215 

1801 

A-267 

3512 

1 A-285 

3932 

A-287 

4033 


» 

Abstracts of Technical Reports on Rockets , Vol. 1: NDRC Publications, as of May 24, 1943, May 1943. 
Heats of Combustion and Formation of Diethylphthalate, Dibutylphthalate , Dinitrotoluene , Diethyl- 
diphenylurea and Nitroguanidine, E. J. Prosen and R. Gilmont, July 1943. 

Abstracts of Technical Reports on Rockets, Vol. 2: Air Corps Jet Propulsion Research Reports, July 
1943. 

Project Summaries for Division 3, Special Projectiles, as of March 1943. 

Interior Ballistics of Recoilless Guns, J. O. Hirschfelder, R. B. Kershner, C. F. Curtiss, and R. E. 
Johnson, Sept. 1943. 

A Microscopical Study of German Powder from a 21 -cm Aircraft Rocket, C. P. Saylor, Apr. 1944. 
Heats of Combustion of Celluloses and Nitrocelluloses , R. S. Jessup and E. J. Prosen, July 1944. 

A Study of the Effect of Radiation on the Burning of Rocket Powder, J. Beek, Jr. 


NONCONTRACT MEMORANDA OF DIVISION 3 


A-1L 


A-4M — A-17M 28 

A-18M — A-27M 29 
A-28M — A-30M 30 


A-38M- 

-A-39M 433 

A-43M 

665 

A-46M 

701 

A-55M 

924 

A-57M 

965 

A-58M 

1022 

A-107M 

4830C 


Derivations of Formulas Used in Computing Effective Gas Velocity and Rocket Velocity from Measured 
Impulse, J. W. M. DuMond, Nov. 1941. 

Notes and Tests of the Design and Performance of Jet Propelled Devices, November 8, 1940 to July 18, 
1941, C. N. Hickman, Sept. 1941. 

Notes on the Design and Performance of Jet Propelled Devices to November 1 , 1941, C. N. Hickman, 
Nov. 1941. 

New Equipment for Measuring and Recording the Pressure of Powder Gases in Rocket Chambers and 
the Thrusts Exerted by Blocked Rockets, Each as a Function of Time, J. W. M. DuMond, Nov. 
1941. 

New Gages for Measuring the Thrusts of Rockets, C. N. Hickman, Feb. 1942. 

A Partial Burning Powder Tester, C. N. Hickman, June 1942. 

Design of Rocket Targets Adopted by the Army Ordnance Department, L. A. Skinner, July 1942. 

Static Tests of the 12-in. J et- Accelerated Armor-Piercing Bomb, C. N. Hickman, Oct. 1942. 

Remarks on the Applications and Performance of Rockets, C. N. Hickman, Oct. 1942. 

Thermal Ignition and Arming Elements for Use with Rockets, C. N. Hickman, Nov. 1942. 

Annotated Bibliography of NDRC Technical Reports and Memorandums of Division 3, Including a 
Listing of Pertinent Reports Issued by Contractors of Division 3, as of September 15, 1944- 


CONTRACT REPORTS FROM DIVISION 3 


Contract No. 
OEMsr-250 

OEMsr-256 

OEMsr-273 

OEMsr-416 

OEMsr-418 

OEMsr-671 

OEMsr-702 


California Institute of Technology 
Included in OEMsr-418. See below. 

Western Electric Company (Bell Telephone Laboratories) 

Many reports, listed below. 

The George Washington University 
Many reports, listed below. 

Hercules Powder Company 

Final Report, July 1, 1943. (Limited distribution to Services.) Development of Jet Propulsion Pow- 
ders, C. W. Gault. 

California Institute of Technology 
Many reports, listed below. 

Budd Induction Heating, Inc. 

Final Report, Sept. 1943. (Limited distribution to Services.) Investigation M-7 Report on Develop- 
ment of 414-in. Rocket under Office of Emergency Management and Scientific Research. Supplement 
to this — Operation Drawings. 

California Institute of Technology 

Final Report, Jan. 1943. (Limited distribution to Services.) Investigations of Double-Base Powders, 
Linus Pauling. (See also later Division 8 reports on Contract OEMsr-881, under which similar 
work was continued at CIT.) 

Interim Reports. (Written under the contract, edited and issued by NDRC.) 


NDRC No. 

OSRD No. 

A-124 

1103 

A-128 

1151 


Investigations of Double-Base Powders: Spectrophotometric Studies, I, R. B. Corey, A. O. Dekker, 
and A. M. Soldate, Dec. 1942. 

X-Ray Diffraction Studies of Molecular Orientation in Double-Base Smokeless Powders Made by the 
Solvent and Solventless Processes, H. A. Levy, Dec. 1942. 




310 


BIBLIOGRAPHY 


NDRC No. OSRD No. 

A-132 1152 Chromatographic Studies of Double-Base Powder I , R. B. Corey, R. Escue, A. L. LeRosen, and 

W. A. Schroeder, Jan. 1943. 

A-151 1265 Measurements of pH on Double-Base Powders , R. B. Corey, C. Green, and H. Levy, Feb. 1943. 

A-194 1558 Investigations of Double-Base Powders , Spectrophotometric Studies , II , R. B. Corey, A. O. Dekker, 

and A. M. Soldate, June 1943. 


Contract No. 
OEMsr-716 


NDRC No. 

OSRD No. 

A- 130 

1188 

A-171 

1370 

A-200 

1713 

A-268 

3544 

A-286 

4009 


Contract No. 
OEMsr-733 


OEMsr-762 


University of Minnesota 

Final Report, October 1945. Studies on Propellants. (In four volumes, including, as appendices, all 
the earlier monthly progress reports.) 

Final Report Supplement. Visual Studies of Propellant Burning. (A two-reel Kodachrome film given 
limited distribution.) 

Interim Reports. (Prepared under the contract, edited and issued by NDRC.) 

The Available Literature on the Mechanism of Combustion of Double-Base Powders , B. L. Crawford, 
Jr., and C. Huggett, Jan. 1943. 

The Ignition by Radiation and Fissuring of Double-Base Powder , B. L. Crawford, Jr., C. Huggett, 
H. S. Isbin, and J. J. McBrady, Apr. 1943. 

Determination of Ignition Temperatures of Double-Base Powders , B. L. Crawford, Jr., and H. S. 
Isbin, July 1943. 

Observations on the Burning of Double-Base Powders , B. L. Crawford, Jr., C. Huggett, and J. J. 
McBrady, Apr. 1944. 

Direct Measurement of Burning Rates by an Electric Timing Method , B. L. Crawford, Jr., and C. 
Huggett, Aug. 1944. 

Duke University 

Final Report, November 1944. Determination of the Linear Burning Rates of Propellants from 
Pressure Measurements in the Closed Chamber, L. G. Bonner. 

Interim Report, June 1944. Title and author as above. 

University of Wisconsin 

Final Report, 1945. (Edited and issued by NDRC.) 


NDRC No. OSRD No. 


A-485 

A-173 

A-243 


6559 

1362 

3206 


Contract No. 
OEMsr-947 


OEMsr-968 


Studies of the Mechanism of Burning of Double-Base Rocket Propellants, Farrington Daniels and 
associates. 

Interim Reports. (Prepared under the contract, edited and issued by NDRC.) 

Testing Powder Grains for Fissures with Special Emphasis on Nonvisual Methods, Farrington 
Daniels and R. E. Wilfong, Apr. 1943. 

The Mechanism of Powder Burning, University of Wisconsin, Jan. 1944. 

Catalyst Research Corporation 

Final Report, Sept. 28, 1943. (Limited distribution to Services.) A Delay Element for Use with 
Rocket Motors, John P. Woolley. 

Budd Wheel Company 
Final Reports, late 1945. 


OSRD No. 

4)132 Rockets for Cast Propellants. 

x/6133 T-59 High Velocity Rocket Grenade (Super Bazooka ). 

u 6134 Rocket Launchers. 

\/6135 Flare Rockets. 

4)136 Airborne Flame Thrower. 

46137 Tank Borne Flame Thrower. 

J6138 Portable Powder Pressurized Flame Thrower. 

46139 115-mm Rocket ( OD-161 , NO-245, Army Int. only). 

4)140 High-Performance Rocket — Design Studies. 
v/6141 Multiple Powder Charge Launcher for JB-2. 

* / 6142 Lightweight 4-2-in. Recoilless Chemical Mortar. 

6143 (This number canceled. The report, on the driver rocket, is included in OSRD 6142.) 
46144 Extension of Range of the 4-2 inch Chemical Mortar “M-.” 

; 6145 Development of a New 4-2 inch Chemical Mortar of Radical Design. 

'4)146 Step Motor Rockets. 

'*4)147 Rocket Powder Traps. 


BIBLIOGRAPHY 


311 


OSRD No. 


^148 
^6149 
^6150 
1^6151 
v-6152 
^6153 
Contract No. 
OEMsr-256 


ItVi inch Rotated Rockets. 

60-mm Recoilless Mortar. 

81 -mm Recoilless Mortar. 

12" Jet Propelled AP Bomb. 

Miscellaneous Development of Rockets and Accessories. 
Summary Report of Rocket Developments. 

Western Electric Company (Bell Telephone Laboratories) 
Final Reports. 


OSRD No. 


i 


6154 Engineering and Material Service , S. R. Avella. 

6155 Mechanical Arming Propeller for 12" Jet- Accelerated AP Bomb , R. F. Mallina. 

6156 Launchers and Improved Components for 4-d-in. Rocket, J. M. Dietz, R. F. Mallina, C. F. Spahn, 

and J. M. Melick. 


6157 

6158 

6159 

6160 
6161 
6162 

6163 

6164 

6165 

6166 

6167 

6168 

6169 

6170 

NDRC No. OSRD 

A-134 1212 

A-187 1484 

A-196 1605 

A-79M ‘ 3035 

A-80M 3082 

A-81M 3083 

A-89M 3446 

Contract No. 


Development of Ribbon Frame Camera, F. L. McNair and F. Reck. 

Ripple Firing Mechanisms for Launching Rockets, D. D. Miller and T. H. Guettich. 

Powder Trapping and the Extrusion of Rocket Powders, R. Burns. 

The Development, Annealing and Calibrating of Copper Tarage Balls, J. R. Townsend. 

The Development of Apparatus for Recording Pressure vs Integral of Pressure, K. S. Dunlap. 

Rocket Launchers for Use on Aircraft, J. M. Dietz, C. A. Hasslacher, and J. H. Mogler. 

Design and Production of Amplifier Calibrators, J. S. Garvin. 

The Design of a 4^2-in. Recoilless Mortar Mount, J. M. Dietz. 

Airborne Flame Thrower Firing Circuits, P. E. Buch. 

The Firing of Rockets by Induction Methods, J. M. Melick. 

Tank Borne Flame Thrower Firing Circuits, P. E. Buch. 

(This number canceled. The report, on an electromagnetic fuze, is included in W-6.1, one of the 
final reports listed below under George Washington University Contract OEMsr-273.) 

X-Ray Diffraction Photograph Investigations of Sheet Powder, W. O. Baker and N. R. Pape. 
Summary Report on Rocket Developments , S. R. Avella. 

Interim Reports. (Written under the contract, edited and issued by NDRC.) 

Basic Flow Properties of Powders of Various Compositions, R. Burns, Jan. 1943. 

Firing Mechanism for the 7 -in. Chemical Rocket Projector, R. F. Mallina and P. T. Higgins, 
May 1943. 

The Ribbon-Frame Camera, F. Reck, July 1943. 

Mechanical Fuze for 12-in. Armor-Piercing Bombs, R. F. Mallina, Dec. 1943. 

Vertical Rocket Launcher for Airplanes, R. F. Mallina, Jan. 1944. 

Jungle Launchers, R. F. Mallina, Jan. 1944. 

Light-Weight Rocket Projectors, R. F. Mallina and J. M. Dietz, April 1944. 


OEMsr-273 The George Washington University (includes Allegany Ballistics Laboratory [ABL]). 

Monograph (unclassified). Mathematical Theory of Rocket Flight, J. B. Rosser and R. R. Newton. 

to be published by the McGraw-Hill Book Company, Inc. 

Final Reports, 1946. 


GWU No. * OSRD No. 

OSRD numbers in the block 5771-5897 which are absent from the list below were canceled. 

B-1.2 5872 Drag of the Propellant Gases on the Powder Charge in Rockets, Part I. Simple Theory, F. T. 

McClure and J. B. Rosser; Part II. Experimental Measurements, J. F. Kincaid and F. T. 
McClure. 


B-2.1 


5861 

5877 


Theory and Application of I e — x dx and I 

Jo Jo 


z , , / y , 

e—V V dy I e — x dx, Part I . Methods of Computation, 


J. B. Rosser; Part II. Applications to the Exterior Ballistic Theory of Rockets , G. L. Gross, 
J. B. Rosser, and E. M. Cook. 


a The GWU numbers group the final reports by type of subject matter, as follows: 
B indicates ballistic theory and design . 

W indicates weapon development. 

P indicates propellant development. 

.1 indicates instrumentation . 


312 


BIBLIOGRAPHY 


GWU No. 

OSRD No 

B-2.2 

5878 b 

B-2.3 

5879° 

B-2.4 

5888 

B-2.5 

5882 

B-2.6 

5883 

B-2.7 

5884 

B-3 

5886 

B-3.1 

5887^ 

B-3. 2 

5889 

B-4 

5890 

B-5 

5891 

B-5.1 

5893 

B-6 

5897 

W-l 

^57 ? 6X 

W-2 

5775 

W-3 

5771 

W-3.1 

5773 

W-3. 3 

5776 

W-3. 4 

5589 

W-4 

5777 

W-5 

5778 d 

W-6 

5779 


W-6.1 

5881 

W-6. 2 

5780 

W-7 

5794 

W-8 

5781 

W-8.1 

5784 

W-8.2 

5785 

W-8. 3 

5786 

W-8. 4 

5788 

W-9 

5789 

W-9.1 

5792 

W-9.2 

5694 

W-9.3 

581 l d 

W-9. 4 

5790 

W-10 

5791 


W-ll 

5800 

W-13.1 

5795 

W-13.2 

5796 


(This is the monograph listed above.) 

Motion of a Spin-Stabilized Rocket during the Burning Period , W. J. Harrington. 

Flight Ballistics Involved in the Use of Rocket-Towed Devices , W. J. Harrington. 

Correlation of Wind Tunnel Data on Rockets , N. G. Gunderson and Seymour Sherman. 

The Relation of Manufacturing Tolerances to Dispersion of Fin-Stabilized Rockets, R. R. Newton 
and M. Goldman. 

Exterior Ballistics of the Cable Bomb, G. L. Gross and J. B. Rosser. 

Some Problems of Heat Transfer in Rockets, J. Beek, Jr., J. B. Rosser, and Harry Siller. 

Temperature Transients in Walls of Rocket Chambers, E. A. Cook and E. H. deButts, Jr. 

(Canceled.) 

Propellant Charge Design of Solid Fuel Rockets, W. H. Avery and J. Beek, Jr. 

The Design of Metal Components for Rocket Motors, H. C. Stumpf and G. W. Engstrom. 

Investigation of Fiberglas Laminates as Materials for Rocket Motors, J. Beek, Jr., and J. F. Kincaid. 

Theoretical Studies of Long-Range and High-Altitude Rockets, J. B. Rosser, F. T. McClure, C. N. 
Hickman, and Nancy Marmer Thompson. 

The J el- Accelerated Armor-Piercing Bomb, C. N. Hickman. 

Rocket Targets Developed for the Army Ordnance Department, C. N. Hickman. 

High-Explosive Anti-Tank 2.36-in. Rocket (Bazooka), C. N. Hickman and S. Golden. 

Interim Ballistic Studies, S. Golden. 

The Development of the T-12 Grenade, D. M. Brasted. 

Development of T4 Powder Charge for M6A3 Rocket Grenade (issued by Division 8), R. Lumry 
and L. Streff. 

Improvement of Components for 4-5-in. Rocket, M8, D. W. Osborne and B. Weissmann. 

Rocket Flares, A. Kossiakoff, N. T. Grisamore, and J. Beek, Jr. 

T-59 High Velocity Rocket Grenade, Part I. Characteristics and Performance, S. Golden; Part II. In- 
ternal Ballistics of Laminated Charges, S. Golden; Part III . Ignition, W. P. Spaulding and L. E. 
Morey; Part IV. Propellant Loss, S. Golden, L. E. Morey, and W. P. Spaulding. 

A Point-Initiating Base-Detonating Electromagnetic Fuze, Allegany Ballistics Laboratory, Bell Tele- 
phone Laboratories, and Explosives Research Laboratory. 

The Follow-Through Rocket Grenade, Tl, W. P. Spaulding and S. Golden. 

Step Motor Rockets, C. N. Hickman and J. M. Woods. 

115-mm Aircraft Rocket, R. E. Gibson and A. Kossiakoff. 

Development of Propellant Charge for 115-mm Aircraft Rocket, J. Beek, Jr., R. L. Arnett, G. W. 
Engstrom, M. Goldman, and A. Kossiakoff. 

Development of Rocket Motor for 115-mm Aircraft Rocket, A. Kossiakoff and G. W. Engstrom. 

Development of Heads and Fuzes for 115-mm Aircraft Rocket, M. J. Walker, F. T. McClure, and 
A. Kossiakoff. 

Development of a High-Performance Composite-Propellant Charge for 115-mm Aircraft Rocket, 
R. Lumry and L. Streff. 

Extension of Range of the 4.2-in. Chemical Mortar M2, G. C. Bowen, C. F. Curtiss, A. R. T. 
Denues, and R. B. Kershner. 

Summary of Interim Ballistic Studies of the 4-3-in. Chemical Mortar, A. R. T. Denues. 

Tests of Various Methods of Obtaining Rotation of the 4-3-in. Chemical Mortar Shell, G. C. Bowen. 

Driver Rocket for the 4-3-in. Chemical Mortar, G. C. Bowen and R. B. Kershner. 

Development of a New 4-3-in. Chemical Mortar of Radical Design, T. R. Paulson. 

Recoilless 4-3-in. Chemical Mortars, Part I. The Development of 4-3-in. Recoilless Chemical Mortar, 
E34R1 , R. B. Kershner; Part II. Corrective Development of 4-3-in. RCM Taken from Final Sum- 
mary Report, No. 12, ABL-CWS 4-3 Mortar Group, A. R. T. Denues; Part III. Development of a 
Cartridge Ignition System, J. M. Woods; Part IV. Driver Rocket Development, S. Golden, J. Levin, 
and F. Culp. 

Propellant Charge Development for 4-5-in. Spinner Rockets, T38E5, T105, T110, D. M. Brasted 
and S. D. Brandwein. 

Rocket for the Anti-Personnel Mine Clearing Snake, Ml, C. A. Boyd and R. J. Bond. 

Investigations of the Use of Rockets to Dispense Mine Clearing Hose, S. D. Brandwein, C. A. Boyd, 
and W. J. Harrington. 


b OSRD number 5878 was assigned to this report (B-2.2), then canceled when its publication as an OSRD monograph was 
approved. 

^Not issued, as of January 1947. Manuscripts loaned by OSRD to Navy Bureau of Ordnance for completion of editing, 
printing, and distribution by Johns Hopkins University Applied Physics Laboratory. 
d Manuscripts not completed, but may be issued under above auspices. 




BIBLIOGRAPHY 


313 


GWU No. 

OSRD No 

W-13.3 

5799 

W-13.4 

5798 

W-13.5 

5801 

W-14 

5802 

W-15 

5803 d 

W-16.1 

5804 

W-16.2 

5805 

W-16.3 

5880 

W-16.3s 

5806 

W-16.4 

5807 

W-17 

5808 

W-18.1 

5812 

W-18.2 

5813 

W-18.3 

5814 

W-18.4 

5787 

W-19 

5815 

W-20 

5818 

W-21 

5793 

W-21.1 

5820 

W-22' 

5821 

W-23. 

5822 


W-24, 

5867 

P-1 

5827 


P-1.1 

5831 

P-1.2 

5833 

P-1.3 

5816 

P-1.4 

5824 

P-2 

5817 


P-2.1 

5832 

P-3 

5828 

P-3.1 

5829 

P-3.2 

5830 

P-4 

5834 

P-5 

5837 

P-6 

5841 

P-6.1 

5842 

P-7 

5844 

P-8 

58,45° 

P-9 

5851 

P-10 

5852 


i 

The Rocket for the Projected Line Charge , C. A. Boyd, W. J. Harrington, and D. Leenov. 

Rocket for Projecting Detonating Cable, C. A. Boyd, D. Leenov, and W. J. Harrington. 

The Rocket for Towing Bangalore Torpedoes, R. J. Bond and C. A. Boyd. 

60-rnm Recoilless Mortar, S. Golden and N. T. Grisamore. 

SI -mm Recoilless Mortar, R. B. Kershner and E. J. Moore. 

(Canceled.) 

Portable One-Shot Flame Thrower, R. E. Hunt. 

Development of Portable Smokeless Powder-Operated Gas Generator for Pressurizing M2-A2 Flame 
Throwers, A. S. Collins and A. A. Nellis. 

Production Model of Portable Smokeless Powder-Operated Gas Generator for Pressurizing M2- A 2 
Flame Throwers (Supplementary Report), Robert Lee James. 

(Canceled.) 

A Gas Generator for a Small Turbine, S. S. Penner and A. J. Madden. 

Booster Launcher for Testing of Aircraft Rockets, M. J. Walker. 

Spiral Launching of 4.6-in. Rockets, R. R. Newton. 

Induction Firing for Rockets, C. F. Bjork and M. Bondy. 

The Development of Rocket Fins and Lug-Band Kits for Use with the Flush-Mount Launcher on 
Aircraft, G. W. Engstrom and R. L. Beddoe. 

The Jet- Assisted Take-Off Unit, L. G. Bonner and W. H. Avery. 

A Multiple-Cartridge Launcher for the JB-2, R. B. Kershner, C. F. Curtiss, V. D. Russillo, and 
C. N. Hickman. 

Design of the High-Velocity Rocket (VICAR), J. Beek, Jr., R. J. Thompson, and R. R. Newton. 

Small-Caliber High-Velocity Rocket (CURATE), R. J. Thompson, D. Brewer, and R. R. Newton. 

The Bumblebee Rocket Motor, S. S. Penner. 

Rocket-Projected Special-Purpose Bombs, Part I. Cable Bomb, A. Africano and J. B. Rosser; 
Part II. Rocket Projection of Incendiary Evaluation Bomb, S. Shulman; Part III . Short-Range 
Rocket-Projected Demolition Bomb, J. F. Kincaid. 

Powder-Driven Post-Hole Digger (Donnerkiel) , R. B. Kershner. 

Burning Rate Studies of Double-Base Powder, W. H. Avery; Part I. Alternate Solventless T2 
Powders, R. E. Hunt and M. N. Donin; Part II. Slow-Burning Powders, R. E. Hunt and 
M. N. Donin; Part III. Temperature Coefficients of Standard Propellants and Promising Experi- 
mental Powders, R. E. Hunt and M. N. Donin; Part IV. Compression Molded Double-Base 
Powders, M. N. Donin. 

Erosive Burning of Double-Base Powders, R. J. Thompson and F. T. McClure. 

Determination of Burning Rates of Certain Powders by the Strand Technique, J. J. Donovan. 

Determination of Burning Rates from Pressure-Time Relations in Closed Chambers, L. G. Bonner. 

Effect of Pressure and Temperature on the Rate of Burning of Double-Base Powders of Different Com- 
positions, W. H. Avery, R. E. Hunt, and L. D. Sachs. 

Studies of Radiation Phenomena in Rockets, Part I . A Study of the Effect of Radiation on the Burning 
of Rocket Powder, J. Beek, Jr.; Part II. Influence of Radiation upon the Burning of Rocket Pro- 
pellants, W. H. Avery; Part III. A Theory of the Effect of Radiation on the Constant Pressure 
Burning Rate of Powders, M. J. Dresher and F. T. McClure; Part IV. Radiation Phenomena 
in Rockets, S. S. Penner. 

Flame Temperature and Radiation Studies in Rockets, R. S. Craig. 

The Reduced Specific Impulse of Ideal Gases, N. Marmer and F. T. McClure. 

A Comparison of the Specific Impulse of Four Double-Base Rocket Propellants, J. P. Rappolt and 
J. Beek, Jr. 

Impulse Determinations of Rockets by Means of Rotating Systems, S. S. Penner. 

Restriction of Powder Burning, A. Turk, L. G. Bonner, A. J. Madden, J. J. Donovan, and W. H. 
Avery. 

A Study of Ignition in the 2.36-in. Rocket Grenade, R. S. Craig and L. D. Sachs. 

Determination of Energies of Explosion of Propellant Powders, J. J. Donovan. 

Certain Special Methods for the Chemical Analyses of Double-Base Powder, J. J. Donovan. 

Dry Extrusion of Powder at Allegany Ballistics Laboratory , H. Higbie and G. F. Padgett. 

Physical Properties of Propellants, H. Higbie. 

Formulation of Manufacturing Specifications for Solid Propellants, R. L. Arnett. 

Miscellaneous Propellant Studies, Part I . Investigation of Some Special Propellant Charge Designs , 
L. G. Bonner; Part II. Utilization of Magnesium as a Rocket Fuel, S. Golden and W. P. Spaul- 
ding. 


314 


BIBLIOGRAPHY 


GWU No. 

P-10.1 

P-10.2 

J-l 

J-l.l 

J-l. 2 

J-1.3 
J-l. 4 
J-l. 5 
J-l .6 
J-l. 7 
J-l. 8 

J-2 
J-2.1 
J-2. 2 


J-3.1 

J-3.2 

J-3.3 

J-4 

J-5 

J-6 


Suppl. No. 
1 
2 

3 

4 

5 

6 


8 

9 

10 

11 

12 

13 

14 

J A-81 
A-93 
A-97 


OSRD No. 

5624 Ballistic Characteristics and Rocket Design Data for Extruded Composite Propellants , R. Lumry and 
L. Streff. 

5853 Captured Enemy Propellants , M. N. Donin and J. J. Donovan. 

5855 Static Range Operational and Fire Control Equipment for Rocket Research , N. E. Alexander and 

C. M. Lathrop. 

5856 The Manufacture of Wire Strain Gages for the Measurement of Pressure as Applied to Rocket Research, 

N. E. Alexander. 

5843 The Manufacture of Wire Strain Gages for the Measurement of Thrust as Applied to Rocket Research, 
N. E. Alexander. 

5857 A-C Bridge and Preamplifier f or Strain-Gage Measurement of Pressure and Thrust, N. E. Alexander. 

5846 Standard Automatic Calibrator for A-C Bridge and Amplifiers, N. E. Alexander. 

5858 Two-Channel Ballistics Camera, N. E. Alexander. 

5869 Standard Frequency Oscillators, Tuning-Fork Type, N. E. Alexander. 

5870 Audible Null Indicator for 15, 000-cycle /sec A-C Bridge Equipment, N. E. Alexander. 

5862 Calibration Equipment for Pressure and Thrust Wire Strain Gages, N. E. Alexander and C. M. 

Lathrop. 

5859 Apparatus for the Recording of Pressure versus Pdt, S. Golden and C. M. Lathrop. 

5847 Electronic Blastmeter, S. Shulman and W. H. Barber. 

5865 Miscellaneous Experimental Electronic Pressure Recorders for Rocket Research, Part I. Electronic 
Sweep Pressure-Time Recorder; Part II. Electronic Maximum Pressure Indicator; and Part III . 
Wide Range Pressures versus Pdt Recorder, N. E. Alexander and W. H. Barber. 

5860 Bourdon System for Pressure Measurement, R. E. Hunt and W. H. Avery. 

5864 Piston Thrust Gage, A. Africano. 

5863 Copper Ball Crusher Gages, A. Africano. 

5868 High-Speed Temperature-Time Recorders Employing an Electronic Inverter, N. E. Alexander. 

5849 X-Ray Photography of Burning Rocket Propellants, S. Golden, R. S. Craig and W. P. Spaulding. 

5871 Application of the Optical-Lever Principle to Following Projection Tip-Off and Yaw Motions During 

Early Burning of Rockets, M. R. Goff. 

Periodical Reports 

Weekly Progress Reports, ABL-WPR 1 (March 30, 1944) through ABL-WPR 63 (Aug. 15, 1945). 
Progress Reports of the Jet Propulsion Research Laboratory, Naval Powder Factory, Indian Head, 
Maryland, 1941 to 1944. 

Supplements to ABL Weekly Progress Reports 
OSRD No. 

3880 Influence of Radiation upon the Burning Rates of Propellants, W. H. Avery, July 8, 1944. 

3960 Preliminary Report on the Motion of a Fin-Stabilized Rocket During the Burning Period, R. R. New- 
ton and J. B. Rosser, Aug. 5, 1944. 

4074 Preliminary Report on the Motion of a Fin-Stabilized Rocket During the Burning Period, Part II, 
R. R. Newton and J. B. Rosser, Aug. 26, 1944. 

4466 Status Report on the T-59 High-Velocity Rocket Grenade, S. Golden, L. E. Morey, and W. P. Spaul- 
ding, Nov. 25, 1944. 

4487 Investigations of a Proposed Liquid Carbon Dioxide Rocket, A. Africano and F. T. McClure, Dec. 2, 
1944. 

4568 Revisions and Corrections to NDRC Formal Report No. A225 , W. H. Avery, R. E. Hunt, and L. D. 

Sachs, Dec. 23, 1944. 

4569 The Rocket Motor for Mine Clearing Snake, M-l; Development and Current Status, C. A. Boyd and 

R. H. Bond, Dec. 23, 1944. 

4694 A Method of Induction Firing for Rockets, C. F. Bjork and M. Bondy, Jan. 20, 1945. 

4942 Effect of Pressure and Temperature on the Rate of Burning of Double-Base Powders of Different Com- 
positions: II, W. H. Avery, R. E. Hunt, and L. D. Sachs, Mar. 2, 1945. 

4905 Tables of Functions Related to the Fresnel Integrals, J. B. Rosser, E. M. Cook, and G. L. Gross, 
Mar. 10, 1945. 

4963 Diffusion of Nitroglycerin in Wrapped Powder Grains, S. Penner and S. Sherman, June 5, 1945. 

4964 Apparatus for the Recording of Pressure vs Pdt, S. Golden and C. Lathrop, June 5, 1945. 

5251 Radiation Phenomena in Rockets, S. S. Penner, Oct. 16, 1945. 

3932 The Status of ABL Projects as of VJ-Day, ABL Staff, Sept. 3, 1945. 

Interim Reports. (Written, in most cases, under the contract, edited and issued by NDRC.) 

793 Tests of the 4-2-in. Chemical Mortar, A. Africano and S. Golden, Aug. 1942. 

888 The Fissuring of Translucent Double-Base Powders at Low Pressures, A. Africano, Sept. 1942. 

929 Pressure Relationships Holding Within Small Rockets, S. Golden, Sept. 1942. 


BIBLIOGRAPHY 


315 


Suppl. No. 

OSRD No 

A-98 

920 

A-107 

1014 


A-113 

992 

A-119 

1079 

A-123 

1100 

A-133 

1226 

A-136 

1142 

A-153 

1303 

A-184 

1460 

A-203 

1658 

A-205 

1703 

A-220 

1886 


A-225 1993 

'I 


A-231 

2085 

A-247 

3232 

A-251 

3280 

,^A-65M 

1156 

A-72M 

1623 

A-75M 

2069 

A-76M 

2069 

A-88M 

3429 


GWU No. OSRDNo. 
ABLsr-1 3711 

2 
2 

(Revised) 


Tests of Cemented Ball Powder Charges , M. Walker and A. Africano; Sept. 1942. 

Thermodynamic Properties of Special Double-Base Powders , D. W. Osborne, F. T. McClure, and 
J. O. Hirschfelder, Oct. 1942. (See also Division 1 Report A-116, OSRD 1087, Thermodynamic 
Properties of Propellant Gases, J. O. Hirschfelder, F. T. McClure, C. F. Curtiss, and D. W. 
Osborne, Nov. 1942. Both came mainly from Division 1 work at the Geophysical Laboratory of 
the Carnegie Institution of Washington under Contract OEMsr-51.) 

Jet Propelled Illuminating Flare, W. E. Jeremiah, Nov. 1942. 

Rockets for Assisted Take-Off of Airplanes ; Their Use and the Prevention of Blast Injury to the Air- 
plane, Leo Maas, Jr., Nov. 1942. 

Dry Extrusion of Double-Base Powder at Indian Head, H. E. Higbie, Dec. 1942. 

Dry Extrusion of Double-Base Powder at Indian Head, II. Extrusion of Solventless Sheet Powder of 
the Russian Formulation, H. Higbie, Jan. 1943. 

Burning Characteristics of Russian Powders, C. F. Bjork and A. Africano, Jan. 1943. 

The Estimation of Pressure-Time Relations Obtaining in Powder Driven Rockets, S. Golden, Mar. 1943. 

Dispersion of a Rotating Rocket, C. H. Dowker, May 1943. 

Extrusions of Double-Base Powder at Indian Head, G. J. Padgett, July 1943. 

Tests of Rotating Budd 4 A in- Rockets, D. W. Osborne and M. J. Walker, Aug. 1943. 

The Extrusion of Dried Solvent Processed Double-Base Powder at Indian Head, H. Higbie, Sept. 1943. 

Effect of Pressure and Temperature on the Rate of Burning of Double-Base Powders of Different Com- 
positions, W. H. Avery and R. E. Hunt, Oct. 1943. 

Relations Between a Rocket and Its Equivalent Shell, J. B. Rosser, Nov. 1943. 

Fin Opening, J. B. Rosser, Feb. 1944. 

Instructions for Use of the Ribbon-Frame Camera, M. J. Walker, Feb. 1944. 

Some Effects of Composition, Powder Temperature and Radiation on the Rate of Burning of Double- 
Base Powders, W. H. Avery, Jan. 1943. 

Changes in the Spider Design in the Extrusion Dies Used at Indian Head, H. E. Higbie, July 1943. 

A Less Regressive Design for Powder Grains, B. Kelly, F. T. McClure, and J. B. Rosser, Nov. 1943. 

Theoretical Study of the Validity of a Certain Method of Determining a Burning Law, B. Kelly, R. B. 
Kershner, F. T. McClure, and J. B. Rosser, Nov. 1943. 

Assembly Operations for Bayonet Igniter Model No. 2, J. W. Burns, Mar. 1944. 

ABL SPECIAL REPORTS 

Rocket Fundamentals, May 1944. (Superseded by ABLsr-4.) 

Technical Conference on Plane to Ground Rocket Tests (Preliminary), Apr. 1944. 

Technical Conference on Plane to Ground Rocket Tests, June 1944. 


3 


3 

(Revised) 

5440 

4 

3992 

5 


6 

5231 

8 

5394 

10 

5548 

11 

5879 

13 

6299 


OSRD No. 


Tables of Ranges and Times of Flight for Certain Values of Angle of Departure, Muzzle Velocity, and 
Ballistic Coefficient, G. L. Gross and W. P. Spaulding, June 1944. 

Tables of Ranges and Times of Flight for Projectiles with Small Ballistic Coefficients, Aug. 1945. 

Rocket Fundamentals, Dec. 1944. 

Description and Facilities of the Allegany Ballistics Laboratory, Cumberland, Md. 

Description and Instructions for the Use of Jet- Assisted Take-Off Unit Model 8AE-1000-H5 , June 1945. 
Preliminary Description of the E37 4 -3-in. Chemical Mortar, Aug. 1945. 

The Reduced Specific Impulse of Ideal Gases, Nancy Marmer and F. T. McClure, Oct. 1945. 
Demonstrations and Descriptions of New 4-3-in. Chemical Mortars of Radical Design, prepared by the 
CWS 4.2 Chemical Mortar Group at ABL, Nov. 1945. 

Program of Events for Press Day at ABL, Nov. 8, 1945, and Memorandum Describing Laboratory 
Facilities and Developments. (Limited distribution to Services.) 

Other reports issued by GWU. (Limited distribution to Services.) 


3974 Photographic Review of Projects Under Investigation at ABL, Review No. 1, Aug. 1944. 

Photographic Review No. 2, Nov. 1944. 

Photographic Review No. 3, Mar. 1945. 

Contents of ABL-WPR Reports. 

Report prepared under the contract, issued by the Chemical Warfare Service 
The 4-3-in. Recoilless Chemical Mortar E34R1- 

Reports prepared at ABL on developments there in which MIT personnel under a Division 11 contract 
collaborated, issued by MIT. 


316 


BIBLIOGRAPHY 


MIT No. 


MIT-MR 151 

MIT-MR 152 

Contract No. 
OEMsr-418 


OSRD No. 
-2544 
^2545 
^2546 
v ^547 
-2548 
/2549 
y2550 

v2551 
v/ 2552 


Development of Portable Cordite Operated Gas Generator for Pressurizing M2-A2 Flame Thrower , A. S. 
Collins and A. A. Nellis, July 1945. 

Powder Pressurizing Unit for Mechanized Flame Thrower , Preliminary Investigations , C. H. King 
and C. A. Boyd, Sept. 1945. 

Powder Pressurizing Unit for Mechanized Flame Thrower Tests with Ignited Fuels , C. H. King, T. Q. 
Eliot, and R. J. Taylor, Sept. 1945. 

California Institute of Technology 

Monographs (unclassified) to be published by the McGraw-Hill Book Company. 

Interior Ballistics of Rockets, Norman Wimpress. 

Exterior Ballistics of Fin-Stabilized and Spin-Stabilized Rockets, Leverett Davis, Jr. and J. W. 
Follin, Jr. 

Manuscript ready (early 1947) for publication under auspices to be arranged. 

Principles of Rocket Design, W. A. Fowler and T. Lauritsen. 

Final Reports, 1946 (Master print sheets of these and of this Summary Technical Report volume are 
to be deposited with the Coordinator of War Department Libraries, for possible photo-offset re- 
production.) 

Ballistic Data, Fin-Stabilized and Spin-Stabilized Rockets. 

Rocket Fuzes. 

Production of Metal Components of Rockets. 

Field Testing of Rockets : Range Operations and Metric Photography . 

Rocket Launchers for Surface Use. 

Firing of Rockets from Aircraft: Launchers, Sights, Flight Tests. 

Aircraft Torpedo Development and Water Entry Ballistics. (Includes list of all CIT reports on this 
subject.) 

Underwater Ballistics and Scale Models of Projectiles. (Includes list of all CIT reports on this subject.) 
Processing of Rocket Propellants . 


In the CIT numbering system, by which the re- 
ports below are listed, the first letter indicates the 
form or use of the report, the second the subject 
matter, and the third the origin (C for California 
on nearly all of those listed). The interim reports 
(“J” series) were widely distributed. Army reports 
in the following list which carry “A” numbers were 
given wide distribution under these numbers rather 
than under the CIT numbers. Of the periodical 
reports, the PMC and LMC series were widely dis- 
tributed . Most of the other reports listed were given 


little or no distribution outside the contract activ- 
ities. Most of the “J” reports were refinements and 
combinations of the “I” reports; the “U” series in- 
cludes many ballistic tables , catalogues of rockets and 
components, bibliographies, abstracts, and indexes. 

Prior to May 1944, OSRD numbers were assigned 
to only a small proportion of the CIT reports, 
mostly those of the “J” series. After that, all “J,” 
LMC, and PMC reports (and a few others) bore 
OSRD numbers in the block 2100-2553, assigned 
exclusively to Contract OEMsr-418. 


INTERIM REPORTS 


CIT No. 

NDRC No. 

OSRD No. 


JAC 1 

A-58 

605 

Controlled Striking Angle of Rocket Projectiles, L. Davis, Jr., and C. F. Robinson, 
Feb. 21, 1942. 26 pp., 3 tables, 5 ill. 

JAC 2 

JAC 3 

A-115 

1069 

Internal Ballistics of Jet-Propelled Devices, B. H. Sage, Oct. 23, 1942. 46 pp., 
5 tables, 27 ill. 

The 6-in. Rocket Motor, J. E. Thomas, Nov. 3, 1942. 10 pp., 2 tables, 7 ill. 

JAC 4 

A-163 

1319 

The Dependence of the Mass of Propellant in a Rocket Motor on the Web Thickness 
and the Motor Dimensions, L. Davis, Jr., and Chester D. Mills, Feb. 25, 1943. 
7 pp., 7 ill. 

JBC 3 

JBC 4 

A-34 

415 

Rocket Targets, W. A. Fowler, Jan. 31, 1942. 35 pp., 4 tables, 11 ill. 

Rocket Targets, W. A. Fowler, Apr. 14, 1942. 22 pp., 6 tables, 5 ill. 

JBC 5x 

A-50 

563 

The Antisubmarine Rocket Projectile, T .Lauritsen, Apr. 25, 1942. 19 pp., 6 tables, 
9 ill. 

JCB 6 

A-57 

585 

The Chemical Warfare Grenade, R. B. King, S. Rubin, and O. C. Wilson, May 20, 
1942. 25 pp., 2 tables, 8 ill. 


BIBLIOGRAPHY 


317 


CIT 

No. 

NDRC No. 

OSRD Ni 

JBC 

7 





A-74I 

758 



A-77II 

803 

JBC 

8 



JBC 

9 



JBC 

10 

A-85 

842 

JBC 

10.2 



JBC 

10.3 



JBC 

10.4 



JBC 

10.5 



JBC 

10.6 



JBC 

11 

A-86 

866 

JBC 

12 



JBC 

13 



JBC 

14 



JBC 

15 



JBC 

16 



JBC 

17 



JBC 

18 



JBC 

19 



JBC 

19.2 



JBC 

20 



JBC 

22 



JBC 

23 



JBC 

24 



Prel. 




JBC 

24 



JBC 

25 



JBC 

26 


2107 

JBC 

27 


2152 

JBC 

28 


2160 

JBC 

29 


2291 

JBC 

30 


2305 

JBC 

31 


2408 

JBC 

32 


2516 


JCC 1 


Antisubmarine Bomb ( ASB ), Parts I and II, W. N. Arnquist and others, June 25, 

1942. 

41 pp., 8 tables, 17 ill. 

45 pp., 10 tables, 19 ill. 

Use of Mousetrap Ammunition , T. Lauritsen, June 27, 1942. 36 pp., 2 tables, 15 ill. 
Use of Subcaliber Mousetrap Ammunition, O. C. Wilson, June 30, 1942. 6 pp., 3 ill. 
Use of 4-5-in. Barrage Rocket, T. Lauritsen, F. C. Lindvall, and L. A. Richards, 
Aug. 1, 1942. 24 pp., 8 tables, 11 ill. 

Use of 4-5-in. Barrage Rocket (Revised), UP Group, Sept. 10, 1942. 35 pp., 6 tables, 
18 ill. 

Installation of BR Projector, UP Group, Sept. 18, 1942. 8 pp., 6 ill. 

Training of Barrage Rocket Crews, W. A. Fowler, Sept. 23, 1942. 11 pp., 2 tables, 

4 ill. 

Training of Barrage Rocket Crews, Comdr. W. F. Royal, W. A. Fowler, and L. A. 

Richards, Oct. 7, 1942. 16 pp., 1 table, 4 ill. 

Manual: Use of 4-5-in. Barrage Rocket — Second Edition, UP Group, Apr. 7, 1943. 
32 pp., 4 tables, 22 ill. 

Chemical Warfare Bomb ( CWB ), R. B. King and W. H. Sleeper, Aug. 20, 1942. 
60 pp., 27 tables, 15 ill. 

The 100-Knot Vertical Flare Mark 4 , J- McMorris, Sept. 21, 1942. 12 pp., 1 table, 
7 ill. 

Mousetrap Operating Instructions, L. B. Slichter and T. Lauritsen, Oct. 25, 1942. 
11 pp., 3 tables, 1 ill. 

CIT Rockets and Test Facilities, an Illustrated Record, Feb. 1, 1943. 64 pp., 1 table, 
59 ill. 

Impact and Deceleration of the ASPC Mark 1 Projectile and Modified A *5 Bomb, 
B. H. Rule and W. P. Huntley, Feb. 10, 1943. 18 pp., 17 ill. 

Installation and Use of Barrage Rocket Projectors for Tank Lighters, May 28, 1943. 

10 pp., 9 ill. 

CIT Rocket Targets, J. B. Edson, June 12, 1943. 33 pp., 5 tables, 31 ill. 
Retro-Bombing: A Description of Projectiles and Installations on Aircraft, June 23, 

1943. 36 pp., 18 tables, 25 ill. 

Ammunition Manual for the 4-5-in. BR ( 1100-yd ), July 26, 1943. 15 pp., 6 tables, 
9 ill. 

Manual for the 4-5-in. Barrage Rocket ( 1100-yd ) — 2nd Edition, Oct. 10, 1943. 19 pp., 
6 tables, 10 ill. 

Ammunition Catalogue: CIT Rockets, Aug. 10, 1943. 151 pp., 54 tables, 99 ill. 
Catalogue: Forward Firing Aircraft Rockets, Nov. 1, 1943. 63 pp., 19 tables, 43 ill. 
Underwater Behavior of 3.5-in. Aircraft Rockets, I. S. Bowen, Dec. 6, 1943. 23 pp., 
6 tables, 5 ill. 

7 .2-in Demolition Rocket — Description and Use, undated. 11 pp., 1 table, 5 ill. 

Manual 7 .2-in. Demolition Rocket, Feb. 10, 1944. 14 pp., 10 ill. 

Brief History of the Development of the 3.5-in. Aircraft Rocket, May 10, 1944. 6 pp., 
3 ill. 

Development of the 3.5-in. Aircraft Rocket, Models 1, 5, and 14, June 1, 1944. 26 pp., 

11 ill. 

Further Investigations of the Underwater Behavior of Aircraft Rockets, I. S. Bowen, 
June 26, 1944. 27 pp., 13 tables, 8 ill. 

Torpedo Deceleration, W. R. Smythe, June 29, 1944. 31 pp., 1 table, 19 ill. 

Manual: Description and Use of the 5.0-in. HVAR, Models 13A and 14A, Nov. 14, 

1944. 31 pp., 3 tables, 21 ill. 

The 2 .25-in. Subcaliber Aircraft Rockets Models 1 and 3, Nov. 20, 1944. 6 pp., 6 tables 

2 ill. 

Preliminary Data, 3.5-in. and 5.0-in. Spin-Stabilized Rockets, Mar. 15, 1945. 31 pp. 

3 tables, 15 ill. 

Land Service Use of 11.75-in. Aircraft Rockets Against Caves, Aug. 15, 1945. 49 pp., 

5 tables, 30 ill. 

Development of Igniter for Cage-Mounted Propellants, J. McMorris and S. Rubin, 
Dec. 19, 1941. 24 pp., 8 tables, 11 ill. 


BIBLIOGRAPHY 


318 

GIT 

No. 

NDRC No. 

OSRD No. 

JCC 

2 

A-56 

597 

JCC 

3 

A-138 

1191 

JCC 

5 

A-158 

1315 

JCC 

6 



JCC 

7 



JCC 

8 



JCC 

9 



JCC 

10 



JCC 

11 



JCC 

12 



JDC 

1 



JDC 

2 

A-104 

947 

JDC 

3.1 

A-35 

445 

JDC 

3.2 

A-39 

473 

JDC 

4 



JDC 

6 



JDC 

8 



JDC 

9 



JDC 

10 



JDC 

11 

A-79 

798 

JDC 

12 



JDC 

13 

A-83 

815 

JDC 

14 

A-84 

818 

JDC 

15 

A-96 

895 

JDC 

16 



JDC 

17 

A-94 

896 

JDC 

18 



JDC 

19 

A- 106 

996 

JDC 

20 

A-56M 

951 

JDC 

21 

A-110 

999 


The Use of Ballistite Turnings in Primers {Preliminary Report), B. H. Sage and W. 

N. Lacey, Mar. 4, 1942. 12 pp., 2 tables, 6 ill. 

A Preliminary Investigation of Plastic Cases for Igniters for Ballistite, B. H. Sage, 
Sept. 15, 1942. 17 pp., 4 tables, 7 ill. 

Investigation of the Use of Plastic-Case Igniters for the ASPC Motor, B. H. Sage, 
Jan. 17, 1943. 25 pp., 14 tables, 14 ill. 

Preliminary Investigation of Metal-Oxidant Igniters for Ballistite, B. H. Sage, Feb.25, 
1943. 20 pp., 7 tables, 13 ill. 

Effect of Relative Humidity on the Water Content of Black Powder, B. H. Sage, May 5, 
1943. 7 pp., 1 table, 6 ill. 

Development of Cellulose Acetate Igniter Cases for 1.25-in. and 2.5-in. Rocket Motors, 
B. H. Sage. 41 pp., 16 tables, 20 ill. 

Effect of Squib Boosters on the Performance of Black Powder Igniters, B. H. Sage, 
Aug. 14, 1943. 18 pp., 9 tables, 1 ill. 

Performance Tests on Electric Squibs and Rocket Igniters After Storage at Elevated 
Temperatures, B. H. Sage, Oct. 16, 1943. 9 pp., 6 tables, 1 ill. 

Threaded-Closure Plastic-Case Igniter for 2.25-in. Rocket Motors, B. H. Sage, Mar. 
16, 1944. 16 pp., 5 tables, 5 ill. 

Development of Tin-Plate Case Igniters for Artillery Rockets, B. H. Sage, Dec. 30, 
1945. 

The Extrusion of UP Propellant, 15/ 16-in. Solventless Extruded Ballistite. T. Lau- 
ritsen, Dec. 15, 1941. 8 pp., 2 ill. 

Some Physical Properties of Ballistite, W. N. Lacey and B. H. Sage, Dec. 27, 1941. 
31 pp., 6 tables, 18 ill. 

Extrusion of Ballistite Tubing and Rod, B. H. Sage and W. N. Lacey, Jan. 20, 1942. 
68 pp., 7 tables, 47 ill. 

Extrusion of Ballistite Tubing and Rod, B. H. Sage and W. N. Lacey, Feb. 23, 1942. 
37 pp., 3 tables, 24 ill. 

Thermodynamic Properties of Products of Reaction of Ballistite, B. H. Sage and W. N. 
Lacey, Feb. 4, 1942. 6 pp. 

Diffusion of Air in Ballistite, B. H. Sage and W. N. Lacey, Feb. 10, 1942. 3 pp., 
1 ill. 

The Temperature of Spontaneous Ignition of Several Samples of American Ballistite, 
P. A. Longwell, B. H. Sage, and W. N. Lacey, Apr. 2, 1942. 7 pp., 3 tables, 2 ill. 
A Study of the Uniformity of the Burning Characteristics of Tubes Extruded from Sol- 
ventless Ballistite, B. H. Sage, D. S. Clark, and W. N. Lacey, Mar. 2, 1942. 14 pp. 
5 tables, 5 ill. 

Charge Design, 2-in. ASB Motor, J. McMorris, June 25, 1942. 27 pp., 1 table, 20 ill. 
Some Effects of Radiation upon Double-Base Powder, B. H. Sage, June 15, 1942. 
39 pp., 7 tables, 20 ill. 

Static Firing Tests on Large-Diameter Grains of Extruded Ballistite, B. H. Sage, 
July 30, 1942. 9 pp., 3 tables, 2 ill. 

Burning Characteristics in the Axial Perforations of Extruded Ballistite Grains, B. H. 
Sage, July 30, 1942. 8 pp., 2 tables, 3 ill. 

Pressure Distribution along Radial-Burning Propellant Grains, B. H. Sage, Aug. 10, 
1942. 18 pp., 9 ill. 

The Influence of Extrusion and Subsequent Storage Upon the Burning Characteristics 
of Ballistite, B. H. Sage, June 1, 1942. 25 pp., 7 tables, 9 ill. 

Influence of Tricresylphosphate upon the Burning Characteristics of Extruded Grains 
of Ballistite, B. H. Sage, July 10, 1942. 8 pp., 4 tables, 2 ill. 

Testing of Quality of Small Grains of Extruded Ballistite, B. H. Sage, Aug. 20, 1942. 
9 pp., 1 table, 3 ill. 

Burning Rate of Four-Spoke Grains of Extruded Ballistite, B. H. Sage, Sept. 25, 
1942. 12 pp., 2 tables, 4 ill. 

Influence of Sizes of the Axial Perforation upon the Performance of Radial-Burning 
Grains, B. H. Sage, Sept. 21, 1942. 11 pp., 1 table, 7 ill. 

Heat Transfer to Nozzles Used in Jet Propulsion Equipment, B. H. Sage, Sept. 30, 
1942. 5 pp., 2 tables, 1 ill. 

Comparative Behaviour of Ballistite from Kenvil and Radford , B. H. Sage, Oct. 14, 
1942. 15 pp., 4 tables, 7 ill. 


BIBLIOGRAPHY 


319 


CIT No 
JDC 24 

JDC 25 

JDC 26 

JDC 28 

JDC 29 

JDC 30 
JDC 31 

JDC 34 

JDC 35 

JDC 36 

JDC 37 

JDC 38 
JDC 39 

JDC 40 

JDC 41 

JDC 42 

JDC 43 

JDC 44 

JDC 45 

JDC 46 

JDC 48 

JDC 49 

JDC 50 

JDC 51 

JDC 52 

JDC 53 

JDC 54 

JDC 55 

JDC 56 

JDC 58 

JDC 59 


NDRC No. OSRD No. 

1183 Extrusion of Large Tubular Grains of Ballistite, B. Hj Sage, Dec. 1, 1942. 46 pp., 

5 tables, 25 ill. 

Effect of Coloring Agents upon the Burning Characteristics of Ballistite , B. H. Sage, 
Dec. 11, 1942. 23 pp., 3 tables, 12 ill. 

1171 Extrusion and Burning Characteristics of a Double-Base Propellant Employing Ethyl 
Centralite as a Stabilizer, B. H. Sage, Nov. 25, 1942. 18 pp., 4 tables, 6 ill. 

1266 Effect of Nitrocellulose Source Upon the Characteristics of Double-Base Powder, B. H. 
Sage, Dec. 15, 1942. 20 pp., 2 tables, 11 ill. 

Extrusion and Burning Characteristics of a Special Propellant, B. H. Sage, Sept. 15, 
1942. 9 pp., 2 tables, 4 ill. 

Some Properties of Solventless Ballistite, B. H. Sage, Dec. 1, 1942. 5 pp., 8 tables. 
Resistance of Ballistite Grains to Internal Pressure, B. H. Sage, Dec. 16, 1942. 7 pp., 

3 tables, 3 ill. 

The Extrusion of Ballistite Dyed with Nigrosine , B. H. Sage, Jan. 21, 1943. 6 pp., 

1 table, 2 ill. 

A-71M 1305 Extrusion and Burning Characteristics of Two Types of Colloidal Propellant, B. H. 

Sage, Jan. 12, 1943. 9 pp., 3 tables. 

Some Studies of the Physical Properties of Ballistite, D. S. Clark, Feb. 11, 1943. 
17 pp., 7 tables, 10 ill. 

Development of a Propellant Grain for Use in a 2-in. Reaction Chamber, B. H. Sage, 
Feb. 10, 1943. 22 pp., 2 tables, 28 ill. 

A-176 1403 Extrusion of Multi-Web Grains of Ballistite, B. H. Sage, Feb. 18, 1943. 13 pp., 16 ill. 

Preparation of Double-Base Propellant for Solventless Extrusions, B. H. Sage, Mar. 1, 
1943.33 pp., 13 tables, 10 ill. 

Extrusion and Burning Characteristics of Three Types of Modified Ballistite, B. H. 

Sage, Mar. 11, 1943. 13 pp., 3 tables, 11 ill. 

Design of a Fast-Burning Propellant Grain for the Barrage Rocket Motor, B. H. Sage, 
Mar. 30, 1943. 12 pp., 2 tables, 11 ill. 

A- 183 1461 Extrusion and Burning Characteristics of Several Modified Ballistites, B. H. Sage, 

Apr. 7, 1943. 15 pp., 3 tables, 9 ill. 

Characteristics of Double-Base Propellants Containing Nigrosine and Carbon Black , 
B. H. Sage, July 16, 1943. 34 pp., 10 tables, 21 ill. 

Design of Dies for the Extrusion of Solventless Ballistite, B. H. Sage, May 29, 1943. 
24 pp., 3 tables, 17 ill. 

Propellant Processing , Igniter Construction, and Motor Loading Facilities as of 
January 1, 1943, B. H. Sage, Nov. 24, 1943. 98 pp., 27 tables, 51 ill. 

Tentative Design of a Cruciform Charge for the 3.25-in. Motor, July 19, 1943. 20 pp., 

4 tables, 15 ill. 

A Study of Certain Hazards Involved in the Loading and Assembly of Rocket Motors, 

A. D. Ayers and B. H. Sage, July 15, 1943. 11 pp., 13 ill. 

An Investigation of the Dispersion of Double-Base Powder in Acetone-Water Mixtures, 

B. H. Sage, Sept. 7, 1943. 11 pp., 6 tables, 4 ill. 

Effect of Storage and Weathering on ASPC Rocket Motors, B. H. Sage, Sept. 23, 1943. 
13 pp., 3 tables, 2 ill. 

Some Physical Properties of Double-Base Powders, B. H. Sage, Oct. 12, 1943. 25 pp., 
8 tables, 15 ill. 

Study of Methods for Evaluation of Quality of Solventless Extruded Ballistite, B. H. 
Sage, Oct. 12, 1943. 5 pp., 4 tables, 5 ill. 

A 12-in. Vertical Press for the Extrusion of Ballistite, B. H. Sage, Oct. 26, 1943. 
10 pp., 2 tables, 30 ill. 

Spontaneous Decomposition of a Ballistite Grain, B. H. Sage, Nov. 20, 1943. 13 pp., 

2 tables, 12 ill. 

Effects of Weathering and Immersion on the Closure Seals of Rocket Motors, B. H. 

Sage, Nov. 20, 1943. 18 pp., 4 tables, 25 ill. 

Development of the Mk 13 Cruciform Propellant Grain, B. H. Sage, Dec. 29, 1943. 

6 pp., 6 tables, 11 ill. 

Corrosive Effect of Solutions of Double-Base Powder on Various Metals, B. H. Sage, 
Jan. 11, 1944. 8 pp., 5 tables, 1 ill. 

Effect of Dimensions on Performance of Tubular Grains, B. H. Sage, Jan. 27, 1944. 

3 pp., 4 tables, 13 ill. 


A-135 

A-137 

A-137 

A-155 


BIBLIOGRAPHY 


320 

CIT No. 

JDC 60 
JDC 61 

NDRC No. OSRD No. 

2134 

JDC 62 

2108 

JDC 64 

2173 

JDC 65 

2205 

JDC 66 

2298 

JDC 67 

2364 

JDC 68 

2365 

JDC 69 

2458 

JDC 70 

2372 

JDC 71 

2348 

JDC 73 

2380 

JDC 74 

2517 

JDC 75 

2541 

JDC 76 
JDC 77 

2519 

2530 

JDC 78 

2532 

JDC 79 

2535 

JDC 80 

2534 

JDC 81 


JDC 82 
JDC 83 
JDC 84 

2537 

2553 

JDC 85 


JDC 86 


JDC 87 


JDC 88 


JDC 89 


JDC 90 


JDC 91 



Vibration Testing of Rocket Motors , B. H. Sage, Feb. 17, 1944. 10 pp., 13 ill. 

Effect of Opacity on the Burning Characteristics of Extruded Ballistite Grains, B. H. 

Sage, Apr. 2, 1944. 20 pp., 9 tables, 15 ill. 

Development of a 24-lb Cruciform Charge for the 5.0-in. Rocket Motor, B. H. Sage, 
May 4, 1944. 13 pp., 3 tables, 13 ill. 

Well Colloided Solvent Process Powder for Solventless Extrusion, B. H. Sage, July 7, 

1944. 23 pp., 8 tables, 16 ill. 

Further Investigation of Partially Colloided, Double-Base Powder in Solventless Ex- 
trusion, B. H. Sage, Aug. 16, 1944. 24 pp., 9 tables, 23 ill. 

The Effect of Processing Operations and Elevated Temperatures upon the Dipheny- 
lamine Content of Ballistite , L. Pauling and B. H. Sage, Sept. 29, 1944. 12 pp., 
5 tables, 5 ill. 

The Investigation of a High-Strength Propellant, B. H. Sage, Nov. 7, 1944. 17 pp., 
12 tables, 4 ill. 

Rate of Diffusion of Nitroglycerin through Cellulose Acetate, B. H. Sage, Jan. 1, 1945. 
14 pp., 5 tables, 6 ill. 

The Measurement of Heats of Explosion and Combustion of Ballistite, B. H. Sage, 
Jan. 13, 1945. 10 pp., 2 tables, 3 ill. 

A Pilot Plant for the Manufacture of Double-Base Propellant by a Modified Solvent 
Process, B. H. Sage, Dec. 1, 1944. 51 pp., 8 tables, 23 ill. 

Ignition Within a Twelve-Inch Vertical Extrusion Press, B. H. Sage, Nov. 24, 1944. 
22 pp., 23 ill. 

Effect of Extrusion Conditions on the Quality of Solventless Ballistite, B. H. Sage, 
Jan. 10, 1945. 34 pp., 5 tables, 7 ill. 

Investigation of JPH Propellant Lots FDAP 28 and FDAP 29, B. H. Sage, June 5, 

1945. 22 pp., 18 tables, 5 ill 

Stabilization of Reaction of Tubular Propellant Grains by the Use of Longitudinal 
Ridges in the Central Perforations, B. H. Sage, May 19, 1945. 28 pp., 6 tables, 
19 ill. 

Intensified Exposure of Motor Seals, B. H. Sage, July 21, 1945. 9 pp., 2 tables, 6 ill. 
Activities of the Technical Supervisors at the Eaton Canyon Pilot Plant, J. I. Gates, 
Oct. 1, 1945. 32 pp., 33 ill. 

Exudation of Nitroglycerin from Ballistite Propellant Grains, B. H. Sage, Sept. 8, 
1945. 13 pp., 8 tables, 4 ill. 

Development of a Triform Grain for 3.25-in. Rocket Motors, B. H. Sage, Sept. 14, 
1945. 35 pp., 11 tables, 19 ill. 

The Compressive Characteristics of Several Propellants Determined at a Constant Rate 
of Stress Application, D. S. Clark, Oct. 11, 1945. 28 pp., 13 tables, 9 ill. 
Compressive, Torsional, and Shear Characteristics of Some Double-Base Propellants , 
B. H. Sage, Nov. 1, 1945. 71 pp., 29 tables, 42 ill. 

Description of Facilities, Eaton Canyon, B. H. Sage, Nov. 1, 1945. 

Description of Facilities at the China Lake Pilot Plant, B. H. Sage, Mar. 13, 1946. 
Investigations on the Burning Characteristics of Propellant Power and Their Effects 
upon Steady-State Pressure in Rocket Motors, B. H. Sage, Nov. 1, 1945. 56 pp., 
9 tables, 32 ill. 

Investigation of Stabilization of Deflagration of Tubular Propellant, B. H. Sage, 
Nov. 15, 1945. 36 pp., 18 ill. 

Impact Characteristics of Several Double-Base Propellants, B. H. Sage, Dec. 10, 1945. 
17 pp., 9 tables, 3 ill. 

The Relation of Column Strength to the Ballistic Performance of Mk 13 Grains, B. H. 

Sage, Dec. 15, 1945. 29 pp., 3 tables, 11 ill. i 

The Effect of Extrusion Pressure on the Degree of Consolidation of JPN Propellant, 
B. H. Sage, Nov. 15, 1945. 

The Use of Ultimate Compressive Strength as Determined in a Simple Compression 
Test as a Measure of JP Propellant Quality for Mk 13 Grains, B. H. Sage, Dec. 
1, 1945. 

Drying Characteristics and the Effect of Water on the Physical Characteristics and 
Ballistic Performance of JPH and JPN Propellants, B. H. Sage, Nov. 15, 1945. 
Ballistic and Physical Characteristics of a Japanese Rocket Propellant, B. H. Sage, 
Dec. 30, 1945. 


BIBLIOGRAPHY 


321 


CIT No. 
JDC 92 

JDC 93 

JDC 94 
JDC 95 

JEC 1 
JEC 2 
JEC 3x 
JEC 4 
JEC 5 

JEC 6 

JEC 6.2 

JEC 6.3 

JEC 7 

JEC 8 
JEC 9 

JEC 9.2 
JEC 9.3 

JEC 10 

JEC 10.2 

JEC 11 
JEC 11.2 


JEC 11.3 
JEC 13 
JEC 13.2 


JEC 14 
JEC 15 

JEC 16 

JEC 17 

JEC 18 
JEC 19 

JEC 20 

JEC 21 

JEC 22 

JEC 23 


NDRC No. 


A-47 

A-50 


OSRD No. 

Free and Restricted Column Behavior of Some Double-Base Propellants, B. H. Sage, 
Dec. 30, 1945. 

Development of a Hexaform Ballistite Propellant Grain for an 8-in. Rocket Motor, 
B. H. Sage, Nov. 15, 1945. 

An Investigation of Nickel-Catalyzed Powder, B. H. Sage, Dec. 15, 1945. 

Preliminary Development Work on the Utilization of Western Cartridge Small Arms 
Powder for Rocket Propellant, B. H. Sage, Dec. 15, 1945. 

A Projector for Target Rockets, W. R. Smythe, Jan. 15, 1942. 12 pp., 9 ill. 

501 A Projector for Target Rockets, W. R. Smythe, Feb. 24, 1942. 17 pp., 9 ill. 

563 The Antisubmarine Rocket Projector, W. R. Smythe, Apr. 27, 1942. 5 pp. 

Projector for the J+Yi-in. Barrage Rocket, July 25, 1942. 11 pp., 10 ill. 

A Twelve-Channel Projector for the Chemical Warfare Bomb (CWB), J.W.M. DuMond, 
June 15, 1942.20 pp., 7 ill. 

Loading of Vertical Bombing Projectors and Preparation of Ammunition { PBY-5 ), 
UP Group, Jan. 6, 1943. 19 pp., 12 ill. 

Crew Manual : Loading of Vertical Bombing Projectors for PBY-5 Aircraft and Pre- 
paration of 205-ft/sec Ammunition, May 1, 1943. 19 pp., 12 ill. 

Crew Manual: Loading of Retro-Bombing Projectors for PBY-5 Aircraft and Pre- 
paration of 200-ft/sec Ammunition: Squadron VP-91, Aug. 14, 1943. 20 pp., 19 ill. 

Crew Manual: Loading of Vertical Bombing Projectors for TBF-1 and TBF-2 
Aircraft and Preparation of 800-ft/sec Ammunition, June 9, 1943. 18 pp., 19 ill. 

Manual: Use of Single-Rail Launcher for 4-5-in. BR, June 26, 1943. 6 pp., 8 ill. 

CIT Type 3 Launcher ( Wooden 3-Rail for 4 -5-in. Barrage Rocket), July 1, 1943. 
20 pp., 20 ill. 

CIT Type 3 Firing Box, July 22, 1943. 12 pp., 8 ill. 

Manual: Instruction for Use of CIT Type 3 Launcher {Wooden 3-Rail for 4 -5 -in. Bar- 
rage Rocket) and CIT Type 3 Firing Box, Nov. 17, 1943. 52 pp., 6 tables, 14 ill. 

Description and Use of Barrage Rocket Launchers {CIT Type 2) for %-ton 4%4 
Truck, July 2, 1943. 13 pp., 13 ill. 

Assembly and Installation of Barrage Rocket Launcher CIT Type 2, for %-ton 4%4 
Truck, July 24, 1943. 10 pp., 8 ill. 

CIT Type 8 Launcher for 4 -5-in. Barrage Rocket, Aug. 4, 1943. 9 pp., 11 ill. 

Manual: Description and Instructions for Use of CIT Type 8 and Type 8 Mod 1 
Launcher for 4-5-in. Barrage Rocket, CIT Launcher Group, revised Nov. 27, 1943. 
50 pp., 10 tables, 47 ill. 

Manual: Description and Instructions for Use, 4-5-in. Rocket Launcher Mk 7 {CIT 
Type 8), Mar. 27, 1944. 60 pp., 6 tables, 41 ill. 

Manual: Description and Instructions for Use, CIT Type 6 Mod 1 Launcher for the 
4.5-in. Barrage Rocket, Nov. 10, 1943. 36 pp., 7 tables, 23 ill. 

Manual: Description and Instructions for Use, CIT Type 6 Mod 1 Launcher for the 
4-5-in. Barrage Rocket {120-Barrel for 2%-ton, 6x6 Amphibious Truck, DUKW), 
Mar. 20, 1944. 26 pp., 19 ill. 

CIT Launcher Catalog, Feb. 7, 1944. 138 pp., 21 tables, 105 ill. 

Manual: CIT Type 7 Mod 1 Launcher for 7.2-in. Rockets, Feb. 4, 1944. 30 pp., 9 
tables, 15 ill. 

CIT Type 9 Launcher {Extensible Single-Rail 4-5-in. BR), May 5, 1944. 31 pp., 
6 tables, 21 ill. 

2167 Manual: Description and Instructions for Use, CIT Type 31C {Shipboard) Launcher, 
H. A. Meneghelli, July 15, 1944. 46 pp., 3 tables, 13 ill. 

2206 Rocket Firing from PT Boats, Sept. 7, 1944. 37 pp., 24 ill. 

2349 Manual: Description and Instructions for Use, CIT Type 44- Launcher for 5.0-in. 
SSR, Jan. 18, 1945. 12 pp., 9 ill. 

2281 Manual: Use of Rocket Launcher Mk 51 Mod 0 {Twelve-Round Automatic for 5.0-in. 
Spin-Stabilized Rockets), Nov. 13, 1944. 31 pp., 10 tables, 22 ill. 

2313 Manual: Description and Use of the 11.75-in. Aircraft Rocket Model 3 from F4U-1D 
Aircraft with Displacement Launcher, Dec. 15, 1944. 43 pp., 4 tables, 29 ill. 

2407 Description and Instructions for Use of Mk 35 Mod 0 Launcher {CIT Type 46B), 
Mar. 15, 1945. 51 pp., 4 tables, 25 ill. 

2422 Description and Instructions for Use, Rocket Launcher Mk 50 Mods 0 and 1 , Apr. 2, 
1945. 74 pp., 2 tables, 26 ill. 


BIBLIOGRAPHY 


322 

CIT No. 
JEC 24 

JEC 25 

JEC 26 
JFC 1 

JFC 2 
JFC 3 

JGC 1 
JGC 2 

JGC 3 

JGC 4 

JGC 5 

JGC 6 

JGC 7 

JGC 8 

JGC 9 

JGC 10 

JHC 1 

JHC 2 

JHC 3 
JHC 4 

JHC 5 

JHC 6 

JHC 7 

JHC 8 

JIC 1 
JIC 2 
JKC 1 

JMC 1 

JMC 2 

JMC 2.2 

JNC 1 

JNC 2 


NDRC No. 


A-62 


A-61 


A-118 


OSRD No. 
2345 

2457 

2448 

632 

2 190 A 
2518 

611 


2484 

2501 

2538 

1082 


2346 

2431 

2533 

2492 

2542 


Rocket Firing from PT Boats , ATB , Ft. Pierce, Florida, P. E. Lloyd, Jan. 13, 1945. 
30 pp., 3 tables, 13 ill. 

Closed-Breech SR Rocket Launcher, CIT Type 38 and 56, Apr. 30, 1945. 31 pp., 
5 tables, 22 ill. 

Rocket Launcher Mk 51 Mod 0, Apr. 25, 1945. 59 pp., 2 tables, 28 ill. 

Photographic Measurements of Rocket Flight, I. S. Bowen and others, June 1, 1942. 
42 pp., 2 tables, 33 ill. 

The Solar Yaw Camera, W. R. Smythe, May 8, 1945. 53 pp., numerous ill. 

The CIT Acceleration Camera Models 3 and 4, submitted by Clyde Chivens, ap- 
proved by I. S. Bowen, Aug. 25, 1945. 19 pp., 1 table, 24 ill. 

Partial Burning of Ballistite Tubes, J. McMorris, Dec. 12, 1941. 7 pp., 1 table, 4 ill. 

Equipment and Procedure for Static Firing Tests, D. S. Clark, Mar. 21, 1942. 24 pp., 
1 table, 15 ill. 

Equipment for Static Firing Tests, E. L. Ellis and N. R. Gunderson, Nov. 12, 1942. 
16 pp., 13 ill. 

Partial Burning Equipment, J. McMorris and F. E. Roach, Nov. 3, 1942. 22 pp., 
20 ill. 

Determination of the Geometrical Malalignment of Rocket Projectiles, T. Lauritsen, 
L. A. Richards, S. Rubin, and J. G. Waugh, Jan. 28, 1943. 14 pp., 1 table, 10 ill. 

Geometic Malalignment in Rockets: Methods of Measurement and Correction; Cor- 
relation with Experimental Results, T. Lauritsen, L. A. Richards, S. Rubin, and 
J. G. Waugh, Aug. 16, 1943. 28 pp., 6 tables, 31 ill. 

Drop Table and Decelerometer: Construction, Calibration and Operation, F. C. Lind- 
vall, Nov. 15, 1943. 12 pp., 7 ill. 

Evaluation of Pressure-Time Relationships Occurring in Static Firing of Rocket 
Motors, B. H. Sage, Apr. 10, 1945. 29 pp., 10 tables, 4 ill. 

Design and Performance of an Installation for the Temperature Conditioning of Motors 
for Static Firing , B. H. Sage, June 1, 1945. 18 pp., 5 tables, 15 ill. 

Some Schlieren Photographs of Rocket Jets, N. U. Mayall, Sept. 25, 1945. 20 pp., 
1 table, 27 ill. 

Test Facilities and Acoustic Range at Morris Dam, B. H. Rule, Oct. 15, 1942. 31 
pp., 24 ill. 

CIT Torpedo Launching Range, Morris Dam Reservoir ( General Description of 
Facilities), F. C. Lindvall, Oct. 15, 1943. 17 pp., 1 table, 12 ill. 

Impact Decelerometer s , B. H. Rule, Jan. 24, 1944. 11 pp., 3 tables, 19 ill. 

Manual: Proof Firing of Rocket Ammunition, 4-5-in. Barrage Rochet, J. D. DeSanto, 
Lt., USNR, Mar. 6, 1944. 29 pp., 1 ill. 

Torpedo Launching Project Report for Year Ending 30 Nov., 1944, F. C. Lindvall, 
Feb. 1, 1945. 56 pp., 65 ill. 

Instruments Developed for .an Experimental Study of the Water Entry of Torpedoes 
and Full-Scale Torpedo Models, F. C. Lindvall, Apr. 10, 1945. 49 pp., numerous 
ill. 

Rocket Testing, U. S. Naval Ordnance Test Station, Inyokern, California, R. W. 
Porter, June 15, 1945. 36 pp., 31 ill. 

Development of a Photo-Flare for Torpedo Launching Investigations, B. H. Sage, 
Dec. 1, 1945. 34 pp., 24 tables, 9 ill. 

Characteristics of PIR Fuze, T. Lauritsen, Aug. 1, 1942. 12 pp., 7 ill. 

Rocket Base Fuzes Mark 162 Mod 0 and Mark 166 Mod 0, July 20, 1945. 15 pp., 6 ill. 

Delay Ejector Units for the 3.5-in. “ Window ” Rocket, B. H. Sage, July 19, 1945. 
36 pp., 5 tables, 25 ill. 

Project Summaries for Division 3, Special Projectiles, as of Sept. 1943. 32 pp., 8 
tables. 

An Introduction to the Study of Rockets, O. C. Wilson and W. N. Lacey, Feb. 5, 1944. 
35 pp., 14 ill. 

An Introduction to the Study of Rockets: Second Edition, O. C. Wilson and W. N. 
Lacey, Mar. 15, 1944. 44 pp., 15 ill. 

Scoring Registers for Target Practice with Automatic Weapons, J. Edson, Feb. 20, 
1942. 16 pp., 5 ill. 

Scoring Register and Recorder for Target Practice, J. Edson, May 1, 1942. 6 pp., 
1 table, 3 ill. 


BIBLIOGRAPHY 


323 


CIT No. NDRC No. 

OSRD No. 


JNC 3 


Feasibility of Visual Coincidence Scoring by Two Observers of Tracer Bullets Shot 
at Rocket Targets, J. W. DuMond, May 1, 1942. 6 pp., 1 ill. 

JNC 4 


Officers’ Manual , Vertical Bombing from PBY-5 Aircraft , Vertical Bombing Section, 
UP Group, Jan. 19, 1943. 35 pp., 8 tables, 23 ill. 

JNC 4 


Supplement to Officers’ Manual Vertical Bombing from PBY-5 Aircraft, Mar. 5, 1943. 

Sup. 


8 pp., 7 ill. 

JNC 4.2 


Officers’ Manual, Vertical Bombing from PBY-5 Aircraft, 205-ft/sec Ammunition, 
May 1, 1943. 37 pp., 8 tables, 25 ill. 

JNC 4.2 


Trajectory Data for 7.2-in. VAR ( 200-ft/sec ), PBY-5, July 15, 1943. 18 pp., 17 ill. 

Sup. 

JNC 4.3 


Officers’ Manual: Vertical Bombing from B-18A Aircraft, 205-ft/sec Ammunition , 
May 15, 1943. 37 pp., 8 tables, 30 ill. 

JNC 4.4 


Officers’ Manual: Retro-Bombing from PBY-5 Aircraft; 200-ft/sec Ammunition; 
Mk 1 Launchers Modified — Squadron VP-91, Aug. 23, 1943. 31 pp., 7 tables, 21 ill. 

JNC 5 


Officers’ Manual: Vertical Bombing from TBF-1 and TBF-2 Aircraft; 300-ft/sec 
Ammunition, June 5, 1943. 38 pp., 7 tables, 30 ill. 

JNC 6 


Use of the 7 -Dial Scoring Register and Tape Recorder for Rocket Target Practice, 
J. Edson, Aug. 11, 1943. 26 pp., 24 ill. 

JNC 7 


Retro-Bombing from B-2/ Aircraft, 300-ft/sec Ammunition, Aug. 1, 1943. 32 pp., 
3 tables, 21 ill. 

JNC 9 


Forward Firing of Rockets from Aircraft, Comdr. J. C. Renard and Lt. Comdr. 
T. F. Pollock, CIT Aircraft Launcher Group, Oct. 4, 1943. 39 pp., 2 tables, 2 ill. 

JNC 9.2 


Forward Firing of 3.5-in. Aircraft Rockets from TBF-1 and PV-1 Aircraft, Nov. 6, 
1943. 79 pp., 14 tables, 45 ill. 

JNC 9.3 


Forward Firing of 3.5-in. and 5.0-in. Aircraft Rockets from TBF-1, PV-1, SBD-5, 
and F6F-3 Aircraft , Dec. 31, 1943. 104 pp., 24 tables, 48 ill. 

JNC 9.3 


Supplement No. 1 to Forward Firing of 3.5-in. and 5.0-in. Aircraft Rockets from 

Sup. 1 


TBF-1, PV-1, SBD-5, and F6F-3 Aircraft, Feb. 26, 1944. 18 pp., 3 tables, 10 ill. 

JNC 10 


Gunnery and Tactical Training with Rocket Targets, J. Edson, Dec. 14, 1943. 87 pp., 
8 tables, 68 ill. 

JNC 10 


Target Launchers (Appendix 2), Dec. 10, 1943. 8 pp., 9 ill. 

App. 

JNC 13 


Handling of Forward-Firing Rocket Equipment Aboard Carriers, Commander Fleet 
Air West Coast, Jan. 8, 1944. 35 pp., 10 tables, 15 ill. 

JNC 13.2 


Handling of Forward-Firing Rocket Equipment Aboard Carriers, Mar. 28, 1944. 
25 pp., 8 ill. 

JNC 14 

2189 

Elements in the Effectiveness of Antisubmarine Attacks by Surface Craft, L. B. Slichter 

Vols. 1 and 2 


and others, May 1944. Vol. 1, 39 pp., 12 tables, 22 ill. Vol. 2, 171 pp., 27 tables, 
73 ill. 

JNC 15 


Theory of Errors in Antisubmarine Attacks by Surface Vessels, N. A. Haskell, Mar. 
16, 1944. 40 pp., 3 tables, 17 ill. 

JNC 16 

2366 

Some Operational and Logistical Problems in the Use of Rockets, W. A. Fowler, Feb. 
1, 1945. 24 pp., 5 tables, 22 ill. 

JNC 22 

2176 

The Effective Temperatures of Rocket Motors with Cruciform Grains, L. Davis, Jr., 
F. E. Roach, and J. M. Schmidt, Aug. 5, 1944. 43 pp., 10 tables, 12 ill. 

JNC 23 

2263 

The CIT Aircraft Rocket Sight Type 2, H. W. Babcock, Oct. 15, 1944. 37 pp., 
4 tables, 20 ill. 

JNC 24 

2333 

Forward Firing of Rockets from P-1/7 D Aircraft, Jan. 2, 1945. 163 pp., numerous 
tables, 39 ill. 

JNC 25 

2382 

Forward Firing of Rockets from P-38L Aircraft, Feb. 17, 1945. 161 pp., numerous 
tables, 34 ill. 

JNC 26 

2347 

Forward Firing of Rockets from P-51K Aircraft, Feb. 10, 1945. 161 pp., numerous 
tables, 35 ill. 

JNC 27 

2476 

Forward Firing of 5.0-in., 3.5-in ., and 2.25-in. Aircraft Rockets from B-25J Aircraft, 
July 15, 1945. 102 pp., 43 tables, 43 ill. 

JNC 28 

2449 

Forward Firing of 5.0-in., 3.5-in., and 2.25-in. Aircraft Rockets from A-26B Aircraft, 
June 15, 1945. 115 pp., 43 tables, 37 ill. 

JNC 29 

2357 

Forward Firing of 11.75-in. Aircraft Rockets from F IffJ and F6F Aircraft (Tentative ) , 
Jan. 27, 1945. 71 pp., numerous tables, 39 ill. 


BIBLIOGRAPHY 


324 


CIT No. 
JNC 29 
Rev. 1 
JNC 29 
Rev. 2 
JNC 29 
Rev. 3 
JNC 30 
JNC 32 

JNC 33 

JNC 34 

JOC 1 

JOC 2 
JOC 3 

JOC 4 

JPC 1 

JPC 2 

JPC 2.2 

JPC 3 

JPC 4 

JPC 5 

JPC 6 

JPC 7 

JPC 8 

JPC 8.2 

JPC 9 

JPC 10 

JPC 11 

JPC 13 
JPC 14 

JPC 15 

JPC 16 
JPC 17 

JPC 18 

JPC 19 


NDRC No. 


A-164 


A-169 

A-73M 


OSRD No. 

Corrections to Sighting Tables , 11.75-in. AR Fired from F4U and F6F Aircraft, 
Apr. 28, 1945. 2 pp. 

Revised Sighting Tables F6F-5 Aircraft 11.75-in. Aircraft Rocket, July 2, 1945. 9 pp., 
7 tables, 2 ill. 

Revised Sighting Tables F4U-1D and F 4 U -4 Aircraft 11.75-in. Aircraft Rocket, Aug. 
15, 1945. 9 pp., 7 tables, 2 ill. 

2428 Principles of Rocket Firing from Aircraft, Illustrated, Apr. 2, 1945. 23 pp., 21 ill. 
2433 Forward Firing of 11.75-in. Aircraft Rockets from A-26B Aircraft, July 1, 1945. 
78 pp., 8 tables, 44 ill. 

2526 Rocket Sight ( Aircraft CIT Type 4), H. W. Babcock, J. L. Fuller, C. A. Domenicali, 
and O. D. Frampton, Aug. 25, 1945. 32 pp., 3 tables, 17 ill. 

2499 Forward Firing of 11.75-in. Aircraft Rockets from P-38L Aircraft, July 20, 1945. 
60 pp., 4 tables, 37 ill. 

Underwater Tests of Depth Charges and Depth Charge Pistols, June 18, 1943. 35 pp., 
14 tables, 24 ill. 

2105 Torpedo Launching Tests at CIT Launching Range through 1 Apr., 1944- 88 pp. 

2381 Structural Damage Associated with Water Entry of Projectiles, D. E. Hudson, Feb. 

15, 1945. 30 pp., 2 tables, 14 ill. 

2539 Development of a Jet-Propulsion Unit for the Mk 13 Torpedo, B. H. Sage, Sept. 12, 
1945. 71 pp., 9 tables, 50 ill. 

Free Flight Trajectories of CWG Projectiles, L. Blitzer, June 1, 1942. 5 pp., 1 table, 
2 ill. 

Curves for External Ballistic Calculations on Low-Velocity Rockets Fired at High 
Angles, L. Davis, Jr., Dec. 17, 1942. 9 pp., 6 ill. 

Curves for External Ballistic Calculations on Low-Velocity Rockets Fired at High 
Angles, Second Edition, L. Davis, Jr., Aug. 20, 1943. 12 pp., 8 ill. 

1330 The Effect of Fin Size, Burning Time, and Projector Length on the Accuracy of Rockets, 

I. S. Bowen, L. Davis, Jr., and L. Blitzer, Jan. 4, 1943. 26 pp., 2 tables, 7 ill. 
Dynamic Stability of Bombs and Projectiles, Chapter I, “Forces On a Solid Moving 

Through an Ideal Fluid,” M. A. Biot, Jan. 2, 1943. 27 pp., 1 table, 6 ill. 
Dynamic Stability of Bombs and Projectiles, Chapter II, “Stability Derivatives in 
a Real Fluid,” M. A. Biot, Jan. 2, 1943. 35 pp., 2 tables, 9 ill. 

Dynamic Stability of Bombs and Projectiles, Chapter III, “Stability of the Rec- 
tilinear Trajectory in Air and Water,” Neglecting Gravity , M. A. Biot, Jan. 2, 
1943. 42 pp., 5 ill. 

1361 Effect of Wind on the Mean Deflection of Rockets, L. Blitzer, Feb. 1, 1943. 8 pp., 
2 tables, 3 ill. 

1632 I: Analysis of the Causes of Dispersion of the 4-3-in. Barrage Rocket: II: Dispersion 
Data on CIT Rockets, May 17, 1943. 13 pp., 7 tables, 2 ill. 

Supplement: I, Analysis of the Causes of Dispersion of the 4-3-in. Barrage Rocket ; 

II, Dispersion Data on CIT Rockets, June 11, 1943. 5 tables. 

Experimental Attempts to Improve the Accuracy of Rockets, O. C. Wilson and G. E. 

Kron, Nov. 26, 1943. 23 pp., 2 tables, 16 ill. 

Dynamic Stability of Bombs and Projectiles, M. A. Biot, Sept. 6, 1943. 123 pp., 
20 ill. 

Dynamic Stability of Bombs and Projectiles, Chapter IV, “Stability of the Vertical 
Fall,” M. A. Biot, May 26, 1943. 25 pp., 1 ill. 

The Mechanism of Water Entry of Projectile, M. A. Biot, Sept. 1, 1943. 38 pp., 5 ill. 
Recorder Settings and Range Errors for Attacks with Ahead-Thrown Weapons on Deep 
Targets, Parts I and II, M. A. Biot, Sept. 1, 1943. 88 pp., 11 tables, 52 ill. 
Comparison of Fin and Rotational Stabilization of Rockets, C. C. Lauritsen, Jan. 
25, 1944. 5 pp. 

Dispersion of Fin-Stabilized Rockets , W. A. Fowler, Jan. 28, 1944. 4 pp. 

Method of Computing Trajectories and Sighting Tables for Forward-Firing Aircraft 
Rockets, L. Blitzer and L. Davis, Jr., Feb. 20, 1944. 33 pp., 15 ill. 

Notes on the External Ballistics of Rotating Rockets, L. Davis, Jr., Apr. 6, 1944. 
26 pp., 1 table, 8 ill 

Sources of Error and Dispersion in Forward Firing of Nonrotating Aircraft Rockets, 
L. Blitzer and L. Davis, Jr., Apr. 25, 1944. 16 pp., 4 ill. 


BIBLIOGRAPHY 


325 


CIT No. 

NDRC No. OSRD No. 

( 

JPC 20 

2188 

Deflection of a Rotating Rocket due to Mallaunching , L. Davis, Jr., Aug. 10, 1944. 
37 pp., 11 ill. 

JPC 20 


Theoretical Curves Showing Yaw, Orientation, and Deflection of 3.5-in. SSR, 5.0-in. 

Sup. 


HCSR, and 5.0-in. HVSR During Burning, L. I. Epstein, Dec. 30, 1944. 9 pp., 

8 ill. 

JPC 21 

2161 

Underwater Trajectories of 3.5-in. Aircraft Rocket Model 5, R. Y. Adams, June 15, 
1944. 15 pp., 2 tables, 9 ill. 

JPC 22 

2235 

Calculation of Mallaunching of Spin-Stabilized Rockets, L. Davis, Jr., and J. G. 
Waugh, Sept. 20, 1944. 21 pp., 4 tables, 7 ill. 

JPC 23 

2190 

Dispersion due to Malalignment of Fin-Stabilized Rockets in Forward Firing from 
Aircraft, L. I. Epstein, Aug. 10, 1944. 9 pp., 6 ill. 

JPC 24 

2528 

The Exterior Ballistics of Fin-Stabilized Aircraft Rockets, L. Blitzer and L. Davis, Jr., 
Aug. 20, 1945. 75 pp., 29 ill. 

JPC 27 

2531 

The Effect of Aerodynamic Moments on the Motion of Spin-Stabilized Rockets during 
Burning, J. W. Follin, Jr., Sept. 21, 1945. 30 pp., 21 ill. 

JPC 29 

2529 

The Influence of the Magnus Moment on the Stability of Rotating Projectiles, L. Davis, 
Jr., and J. W. Follin, Jr., Sept. 1, 1945. 19 pp., 6 ill. 

JPC 30 

2527 

A Theory for the Difference Between the True Angle of Attack and the Effective Angle 
of Attack, L. Davis, Jr., Sept. 28, 1945. 9 pp., 2 ill. 

JPC 31 

2536 

Range Tables for Spin-Stabilized Rockets, J. W. Follin, Jr., and P. W. Stoner, 
Nov. 15, 1945. 73 pp., 53 tables. 

JQC 1 

2525 

Rocket Terminal Ballistics Facilities at NOTS, Inyokern , and Test Results to 15, 
April 1945, J. E. Thomas, Aug. 10, 1945. 61 pp., 3 tables, 43 ill. 

JSC 1 


Manual: Methods of Manufacture for the 4-5-in. BR, Lowell Martin, Sept. 22, 1943. 
89 pp., 2 tables, 80 ill. 

JSC 2 


Manufacturing Methods for 3.5-in. Rocket Motor Mk 7 and 3.5-in. Rocket Body Mk 1 
CIT Section “B,” Jan. 20, 1944. 55 pp., 4 tables, 42 ill. 

JSC 3 


Manual: Manufacturing and Inspection Problems: Rocket Target Mk 3, Develop- 
mental Engineering Section, Mar. 1, 1944. 32 pp., 18 ill. 

JSC 4 


Methods for One-Piece Nozzle Manufacture, Developmental Engineering Section, 
Mar. 8, 1944. 51 pp., 31 ill. 

JSC 4 


Supplement No. 1: Methods for One-Piece Nozzle Manufacture, Developmental 

Sup. 1 


Engineering Section, May 16, 1944. 3 pp., 3 ill. 

JSC 5 

2110 

Manual: Inspection Procedures for 3.5-in. Aircraft Rocket Model 5 { 3.25-in . Rocket 
Motor Mk 7 and 3.5-in. Rocket Body Mk 1), Developmental Engineering Section, 
Apr. 29, 1944. 33 pp., 20 ill. 

JSC 6 

2119 

Manual: Inspection Procedures for 2.25-in. Aircraft Rocket Model 3 { Subcaliber ), 

( 2.25-in . Rocket Motor Mk 12 and 2.25-in. Rocket Body Mk 1), Developmental 
Engineering Section, May 30, 1944. 28 pp., 19 ill. 

JSC 7 

2204 

Manual: Inspection Procedures for 5.0-in. High-Velocity Aircraft Rocket: Models 
13, 14, 15, and 16. {5.0-in. Rocket Motor Mk 1 and 5.0-in. Rocket Body Mk 5), 
Developmental Engineering Section, Aug. 28, 1944. 40 pp., 27 ill. 

JSC 7 

2234 

Manual: Inspection Procedures for 5.0-in. High-Velocity Aircraft Rocket, Supple- 

Sup. 1 


ment No. 1. Inspection of 5.0-in. Rocket Body Mk 5 Mod 1 , Developmental Engi- 
neering Section, Sept. 1944. 12 pp., 9 ill. 

JSC 8 

2236 

Instruction Manual: Optical Inspection Fixture, Model M-4 , 5.0-in. High-Velocity 
Aircraft Rocket, Developmental Engineering Section, Oct. 2, 1944. 23 pp., 11 ill. 

JSC 9 

2311 

Manual: Inspection Procedures for 11.75-in. Rocket Motor Mk 1, Nov. 22, 1944. 
42 pp., 20 ill. 

JSC 10 

2383 

Manual: Inspection Procedures for 5.0-in. /5 High-Capacity Spin-Stabilized Rocket 
CIT Model 34 {5.0-in. Rocket Motor Mk 4 cmd 5.0-in. Rocket Head Mk 10), De- 
velopmental Engineering Section, Jan. 15, 1945. 45 pp., 26 ill. 

JSC 10 

2421 

Manual: Inspection Procedures for 5.0-in. Spin-Stabilized Rockets. Supplement 

Sup. 1 


No. 1. Inspection of 5.0-in. Rocket Head Mk 8 {for 5.0-in./ 10 Common Spin- 
Stabilized Rocket, CIT Model 32), Developmental Engineering Section, Mar. 15, 
1945. 15 pp., 13 ill. 

JTC 1 

2284 

Effects of Rocket Blast on Aircraft Structures, E. C. Briggs and C. H. Wilts, Nov. 
16, 1944. 36 pp., 1 table, 17 ill. 


326 


BIBLIOGRAPHY 


CIT No. NDRC No. OSRD No. 

PERIODICAL REPORTS 


LMC 1.1 to 
1.26 



UMC 

UMC 

UMC 


1.1 to 
1.86 
1 .44 to 
1.180 

1.1 to 

I. 99, 

2.1 to 
2.89 

6. lx to 
6.57x 

11.1 to 

II. 14 
45. lx to 
45.14x 


Confidential Bulletin — Section L. Division 3, NDRC, published semi-monthly by 
OSRD Contract OEMsr-418, CIT, May 15, 1944 to July 15, 1945. (LMC 1.1 to 
1 .4 were called the Newsletter, but the name was changed to Confidential Bulletin 
with the issue of LMC 1.5, Aug. 1, 1944.) 

Local Progress Reports, Oct. 26, 1941 to Jan. 20, 1943. (Discontinued.) 

Morris Dam Weekly Reports (Biweekly beginning Apr. 1, 1945), Aug. 23, 1942 to 
Aug. 4, 1945. 

Weekly Progress Reports (for Washington), Oct. 26, 1941 to Sept. 15, 1945. (Bi- 
weekly, beginning Apr. 1, 1945.) 


Abstracts of British Acquisitions, July 24, 1943 to Aug. 31, 1945. 

Bibliography of Published Reports, Aug. 1, 1943 to Aug. 1945. (Appeared monthly.) 

Annotated Bibliography of Reports Received from Outside Sources, Jan. 1, 1944 
through July 22, 1944, July 24, 1944 through Aug. 19, 1944, Aug. 21, 1944 
through Sept. 30, 1944, thereafter monthly for Oct. 1944 through Aug. 1945. 


CZC 1 
DDC 1 

IAC 1 
IAC 2 
IAC 3 

IAC 4 
IAC 5 

IAC 6 

IAC 7 

IAC 8 

IAC 9 
IAC 11 
IAC 12 

IAC 13 

IAC 14 

IAC 15 
IAC 16 

IAC 17 

IAC 18 

IAC 19 

IBC 1 
IBC 2 
IBC 3 


OTHER REPORTS PUBLISHED 

2395 Explosives Safety Regulations, Feb. 15, 1945. 11 pp., 1 table. 

Resume of Visit to the Sunflower Ordnance Works, Lawrence, Kansas, W. H. Cor- 
coran, Nov. 2, 1944. 25 pp., 3 tables. 

Refractory Linings for UP Motors, J. McMorris, Feb. 13, 1942. 2 pp. 

High Performance 2-in. APR, F. E. Roach, Aug. 13, 1942. 7 pp., 3 tables, 3 ill. 
Motors for Antisubmarine Bombs, Vertical Bombs and Barrage Rockets, W. A. Fowler, 
Oct. 22, 1942. 6 pp., 2 tables. 

Design of Box Grids, S. Rubin, Nov. 6, 1942. 3 pp., 1 ill. 

Some Factors Entering into the Design of High Performance Rockets, E. Ellis and F. 

Roach, Jan. 10, 1943. 12 pp., 2 tables, 2 ill. 

Influence of Burning Time, Mass Velocity, and Tube Wall Thickness on the Heat 
Failure of Rocket Tubes, Motor Test Section, Jan. 20, 1043. 23 pp., 5 tables, 15 ill. 
Standardization of 3.25-in. OD Motors Using 3.5 lb of Propellant, Motor Test 
Section, March 8, 1943. 16 pp., 4 tables, 10 ill. 

Influence of Temperature upon Design and Application of Rocket Motors, May 4, 
1943. 7 pp., 1 table, 3 ill. 

Plastic Subcaliber N ozzle Tests, J. A. Gilbert, June 9, 1943. 4 pp., 1 table, 1 ill. 

Tests of Double-Ended Rocket Motors at Goldstone, Aug. 5, 1943. 6 pp., 2 tables. 
Proposed Design of Rocket Motor for Mk 13 Torpedo, B. H. Sage, Nov. 15, 1943. 
27 pp., 12 ill. 

Nozzle Erosion as a Function of the Physical Properties of the Material, June 22, 1944. 
9 pp., 2 tables, 3 ill. 

Nozzle Erosion in the 3MR3 Rocket Motor Determined from Static Firing Records, 
N. U. Mayall, Aug. 7, 1944. 11 pp., 8 ill. 

5.0-in. Rocket Motor ( 5MR5 ), M. C. Pond, Dec. 19, 1944. 37 pp., 21 ill. 
Specifications for Standard Assembly of 5.0-in. Rocket Motor Mk 6 Mod 2, M. C. 
Pond, May 15, 1945. 44 pp., 22 ill. 

Specifications for Standard Assembly of 5.0-in. Rocket Motor CIT Model 38 (5MA5), 
M. C. Pond, May 29, 1945. 56 pp., 1 table, 34 ill. 

Investigation of Several Types of German Rocket Motors, B. H. Sage, Aug. 25, 1945. 
31 pp., 6 tables, 23 ill. 

Closures and Seals for Rocket Motors, B. H. Sage, Nov. 19, 1945. 52 pp., 11 tables, 
19 ill. 

ATG Report, S. Rubin, Jan. 13, 1942. 2 pp. 

Chemical Warfare Grenade ( CWG ), S. Rubin, Feb. 5, 1942. 5 pp., 1 table. 1 ill. 

A$ Projectiles, V. Rasmussen, W. N. Arnquist, V. Anthony, and T. Lauritsen, 
Apr. 11, 1942.3 pp. 


NDRC No. OSRD No. 


BIBLIOGRAPHY 


327 


CIT No. 
IBC 4 

IBC 5 
IBC 6 
IBC 7 
IBC 8 

IBC 9 
IBC 10 

IBC 11 

IBC 12 

IBC 13 

IBC 14 

IBC 15 

IBC 17 

IBC 19 

IBC 20 
IBC 21 
IBC 22 

IBC 22.2 
IBC 22.3 
IBC 23 

IBC 24 

IBC 26 
IBC 26.2 

IBC 27 
IBC 28 
IBC 29 
IBC 29.2 

IBC 30 

IBC 31 

IBC 32 
IBC 33 
IBC 34 
IBC 34.2 

IBC 35 
IBC 36 
IBC 37 
IBC 38 

IBC 39 

IBC 40 


Development of Separable Motor for AS Model Bomb Tests, W. N. Arnquist, Apr. 28, 

1942. 3 pp., 1 table. 

Status of High Altitude Rocket Work at CIT, W. A. Fowler, June 8, 1942. 5 pp. 
Motor Design, F. E. Roach, Aug. 10, 1942. 9 pp., 6 ill. 

Computations on Poppet Valve for ASB Motor, S. Rubin, Oct. 22, 1942. 2 pp. 

Entry and Underwater Characteristics of 7 . 2-in. -Diameter , Flat-Nosed Mousetrap and 
Hedgehog Projectile, B. H. Rule, July 13, 1942. 32 pp., 4 tables, 17 ill. 

The ASB Retrieving Line, J. Edson, June 3, 1942. 5 pp. 

Deflection and Dispersion of ASB’s, C. W. Snyder, Nov. 25, 1942. 8 pp., 5 tables, 

1 ill. 

Single Shroud Rocket Tail with Internal Insulated Firing Ring, L. A. Richards, 
Jan. 22, 1943. 4 pp., 3 ill. 

Six-Inch Scatter Bomb, Tests of Preliminary Design Shapes, B. H. Rule and W. P. 
Huntley, Jan. 12, 1943. 5 pp. 

The Effect of Burning Time, Fin Size and Projector Length on the Accuracy of the 
1800-ft/sec 2-in. A A, L. Blitzer, Feb. 8, 1943. 7 pp., 5 tables, 2 ill. 

Tests of Lateral Dispersion of BR with Various Nozzles, L. A. Richards and J. G. 

Waugh, Nov. 24, 1942 to Jan. 14, 1943. 11 pp., 8 tables. 

Vertical Bombing with ASB Std., B. H. Rule and W. P. Huntley, Jan. 1943. 9 pp., 
4 tables, 4 ill. 

Tests of Fast Burning BR with AIR 2 Fuse, 25-26 Feb. 43, C. Snyder and V. Ras- 
mussen. 13 pp., 4 tables, 5 ill. 

Tests of Lateral Dispersion of BR with Various Nozzles, J. G. Waugh and L. A. 

Richards, Jan. 30 to Feb. 5, 1943. 14 pp., 7 tables, 4 ill. 

BR Firing of 2 Mar 43, at MAAR, Accuracy Committee. 12 pp., 1 table, 6 ill. 
Torpedo Deceleration, W. R. Smythe, Apr. 10, 1943. 23 pp., 15 ill. 

Status of 7V11; Preliminary Standardization Tests, Projectile Section, Apr. 14, 1943. 
9 pp., 5 tables. 

Field Standardization of the 7V 11 , May 24, 1943. 16 pp., 14 tables, 2 ill. 

Revised Field Standardization of 7 . 2-in. VAR, June 1943. 5 pp., 4 tables, 2 ill. 
Gas-Malalignment and Deflection-Malalignment Ratio for All Types of BR Fired 
from September 20, 1942 to April 1, 1943, C. W. Snyder. 15 pp., 2 tables, 8 ill. 
Mechanical Destruction Tests of Metal Parts of Subcaliber Mk 1 ASPC Tails, B. H. 

Sage, Apr. 20, 1943. 14 pp., 1 table, 4 ill. 

Field Standardization of the 7V13, May 12, 1943. 13 pp., 12 tables, 2 ill. 

Revised Field Standardization of 7.2-in. VAR ( 400 ft/sec), July 19, 1943. 5 pp., 
4 tables, 2 ill. 

Description of Some CIT Projectiles, May 19, 1943. 19 pp., 13 ill. 

Barrage Rocket Packing Box Projector, May 22, 1943. 6 pp., 1 table, 4 ill. 

Field Standardization of the 7V12, May 24, 1943. 13 pp., 11 tables, 2 ill. 

Revised Field Standardization of 7.2-in. VAR (300 ft /sec), July 19, 1943. 5 pp., 
4 tables, 2 ill. 

Preliminary Field Standardization of the IRG, J. McMorris, June 18, 1943. 16 pp., 
3 tables, 11 ill. 

Comparative Trajectories of Various A A Rockets, L. Davis, Jr., June 3, 1943. 4 pp., 

2 ill. 

Status Report on Smoke Float Rockets, S. Rubin. June 7, 1943. 4 pp., 2 ill. 

Field Standardization of ASPC Subcaliber, June 24, 1943. 5 pp., 3 tables, 2 ill. 

Field Standardization of 300-ft/sec Subcaliber VAR, June 1943. 5 pp., 3 tables, 2 ill. 
Field Standardization of the 2.5-in. Rocket Mk 3 Mod 1 (300-ft/sec Subcaliber), Pro- 
jectile Section, Oct. 18, 1943. 5 pp., 3 tables, 2 ill. 

Field Standardization of 200-ft/sec VAR Subcaliber, June 1943. 5 pp., 3 tables, 2 ill. 
Range Standardization of 4-5-in. Barrage Rocket, July 9, 1943. 11 pp., 1 table, 10 ill. 
Field Standardization of 7.2-in. ASR, July 11, 1943. 10 pp., 3 tables, 7 ill. 

Static and Field Standardization of the Short-Burning 250-yd 4-3-in. BR, July 31, 

1943. 6 pp., 4 tables, 2 ill. 

Accuracy of the CWR-N, C. Weinland, J. W. McConnell, and F. W. Thiele, Sept. 
25, 1943. 16 pp., 11 tables, 6 ill. 

Static and Field Standardization of the Short-Burning 1000-yd 4-3-in. BR, Projectile 
Section, Aug. 8, 1943. 14 pp., 8 tables, 5 ill. 

X 


328 


BIBLIOGRAPHY 


CIT No. 
IBC 41 

IBC 42 

IBC 43 

IBC 46 

IBC 46.2 
IBC 47 

IBC 48 

IBC 49 

IBC 50 
IBC 51 

IBC 52 

IBC 53 

IBC 54 
IBC 56.1 

IBC 56.2 

IBC 56.3 

IBC 57 
IBC 58 

IBC 59 

IBC 60 

IBC 61 

IBC 62 

IBC 64 

IBC 66 

IBC 67 
IBC 68 

IBC 69 

IBC 70 

IBC 71 
IBC 72 

IBC 73 


NDRC No. OSRD No. 

Field Standardization of the VFR — CIT 200-ft/sec, Aug. 1, 1943. 4 pp., 2 tables, 
2 ill. 

External Ballistics of the 2.5-in. Rocket Grenade ( 300-ft/sec ) Based upon Data for 
the 2.5-in. RG {350-ft /sec) , July 29, 1943. 9 pp., 4 tables, 6 ill. 

Field Tests of the 3.25-in. Rocket Projectile 3A9, May 19 to June 28, 1943. 9 pp., 
2 tables, 2 ill. 

Preliminary Standardization of the CWR-N, Projectile Section, Sept. 3, 1943. 4 pp., 

2 tables, 2 ill. 

Standardization of the CWR-N, Projectile Section, Oct. 12, 1943. 6 pp. 

Comparison of Design and Performance of 7-in. CWR and 7.2-in. CWR-N , T. Laurit- 
sen, Sept. 14, 1943. 5 pp., 6 tables. 

Preliminary Standardization of the 3.5-in. AR { 3A12 ), Projectile Section, Sept. 18, 
1943 . 8 pp., 4 tables, 2 ill. 

The Cause of Abnormally Short Range BR Shots, J. E. Thomas, Oct. 1, 1943. 3 pp., 
1 table. 

Notes on Target Fin Construction, W. D. Lacey, Oct. 9, 1943. 2 pp. 

Comparison between the 81 -mm Mortar and the 4-5-in. BR, E. Thomas and P. Lloyd, 
Sept. 23, 1943. 3 pp. 

Static and Field Standardization of the 3.5-in. AR Model 1 , Projectile Section, Oct. 9, 

1943. 7 pp., 3 tables, 2 ill. 

BR Parachute Drops, J. E. Thomas and P. E. Lloyd, Oct. 20, 1943. 8 pp., 1 table, 
10 ill. 

Smoke Float Rocket, S. Rubin, Nov. 24, 1943. 3 pp. 

Preliminary Tests on 7 .2-in. -Diameter UWR Assembly, B. H. Rule, Nov. 19, 1943. 

3 pp., Morris Dam Report No. 91. 

U.W.R. Projectile, B. H. Rule and W. P. Huntley, Jan. 10, 1944. 33 pp., 3 tables, 
22 ill., Morris Dam Report No. 91. 

U.W.R. Projectile, B. H. Rule and W. P. Huntley, Apr. 19, 1944. 8 pp., 4 tables, 
Morris Dam Report No. 91, Addendum 1. 

BR Parachute Drops, P. E. Lloyd and R. D. Ridgeway, Dec. 6, 1943. 16 pp., 18 ill. 
Static and Field Standardization of the 3.5-in. AR Model 5, Projectile Section, Jan. 
5, 1944. 9 pp., 4 tables, 5 ill. 

Static and Field Standardization of the 5.0-in. AR Model 107, Projectile Section, 
Jan. 5, 1944. 7 pp., 4 tables, 4 ill. 

Underwater Performance Tests of 7.2-in. Rocket Mark 3 Mousetrap Assembly with 
Mark 131 Fuze and with Mark 140 Fuze and Protective Cap, R. L. Noland and 
B. H. Rule, Feb. 4, 1944. 7 pp., 1 table, 4 ill. Morris Dam Test No. 101. 

Static and Field Standardization of the 7.2-in. Demolition Rocket, Projectile Section, 
Feb. 1, 1944. 6 pp., 3 tables, 3 graphs. 

Dispersion and High Temperature Limit of the CWR-N , Projectile Section, Feb. 15, 

1944. 5 pp., 4 tables, 2 ill. 

Status of the 5.0-in. HVAR {5.0-in. Motor) as of 1 April, 1944, C. W. Snyder. 8 pp., 
3 tables, 5 ill. 

Static Tests of the 3MR12 Motor Mounted for Rotation, N. U. Mayall, May 28, 1944. 
19 pp., 4 tables, 10 ill. 

Status of 12.0-in. Aircraft Rocket, S. Rubin, June 2, 1944. 14 pp., 16 ill. 
Preliminary: Assembly Instructions, Safety Precautions, and Firing Instructions 
for 5.0-in. High-Velocity Aircraft Rocket, June 23, 1944. 10 pp., 7 ill. 

U. S. Navy 5.0-in. High-Velocity Aircraft Rocket {HVAR) Description and Per- 
formance, June 23, 1944. 15 pp., 10 ill. 

2174 Preliminary: Sighting and Trajectory Drop Tables for 2.25-in. Aircraft Rockets, 
Angle of Attack Tables for F4U-1, F6F-3, SBD-5, SBD-6, TBM-1 , TBF-1C, 
Aug. 5, 1944. 73 pp., 72 tables. 

Trajectory Drops for 5.0-in. HVAR {JPN), Aug. 24, 1944. 8 pp. of tables. 
Preliminary: Excerpts from CIT Complete Projectile Catalog, Aug. 25, 1944. 35 pp., 
12 tables, 21 ill. 

Range of the 5.0-in. AR {JPN) Model 5 and Model 11 Rockets when Ground Fired 
from the Aircraft Rocket Launcher Mk 4 } J- G. Waugh and J. N. McClelland, 
Jan. 27, 1945. 6 pp., 1 table, 3 ill. 


NDRC No. OSRD No. 


BIBLIOGRAPHY 


329 


CIT No 
IBC 74 

IBC 75 

ICC 1 

ICC 2 

ICC 3 

IDC 1 
IDC 2 
IDC 3 

IDC 4 

IDC 6 

IDC 7 
IDC 9 

IDC 10 

IDC 11 
IDC 12 
IDC 13 

IDC 14 

IDC 15 

IDC 16 

IDC 16 
Rev. 
IDC 17 

IDC 18 

IDC 19 
IDC 20 

IDC 21 
IDC 22 
IDC 23 

IDC 24 
IDC 25 
IDC 26 


Note on Range and Dispersion of 5.0-in. HVAR Model 13, H. M. Greene and L. 

Davis, Jr., Mar. 27, 1945. 5 pp., 1 table, 2 ill. 

Design and Development of the 11.75-in. Rocket Motor, C. W. Snyder, Nov. 6, 1945. 
43 pp.,28 ill. 

Description of an Igniter for Mousetrap Propellant, Aug. 10, 1942. 11 pp., 1 table, 

5 ill. 

A Preliminary Investigation of Plastic Cases for Igniters for Ballistite, B. H. Sage, 
Sept. 15, 1942. 25 pp., 5 tables, 9 ill. 

Development of a Toroid Igniter for Application in the 3.25-in. Spin-Stabilized Rocket 
Motor Mk 13, B. H. Sage, Nov. 15, 1945. 9 pp., 4 tables, 1 ill. 

Extrusion of UP Propellant, T. Lauritsen, Dec. 11, 1941. 9 pp., 1 ill. 

Partial Burning of DBP Sticks, J. McMorris, Dec. 12, 1941. 8 pp., 3 tables, 4 ill. 
Pressure Build-Up in Partial-Burning Tests, J. McMorris and W. N. Arnquist, 
Jan. 21, 1942. 3 pp., 1 table, 1 ill. 

Burning Properties of Solventless Extruded Propellant: Preliminary Results, B. H. 

Sage and W. N. Lacey, Jan. 27, 1942. 9 pp., 5 ill. 

A Study of the Uniformity of Deflagrating Characteristics of Extruded Solventless Sheet 
Ballistite , B. H. Sage and W. N. Lacey, Feb. 20, 1942. 20 pp., 3 tables, 6 ill. 
Cemented Joints of Ballistite Columns, J. McMorris, Feb. 13, 1942. 1 p. 

Effect of Added Coloring Agents Upon the Burning Characteristics of Ballistite, B. H. 

Sage, Oct. 1, 1942. 28 pp., 3 tables, 14 ill. 

Some Calculations and Experimental Measurements Upon the Pressure Distribution 
Around Thin-Webbed Charges During Firing, R. N. Wimpress, G. W. Miller, 
B. H. Sage, and W. N. Lacey, Apr. 8, 1942. 20 pp., 2 tables, 9 ill. 

The Effect of Temperature on the Behavior of the CWG, J. McMorris, Apr. 17, 1942. 

6 pp., 2 tables, 3 ill. 

Internal Ballistics of the 1 Y%-in. Motor Using l]/i-in. and 1 x / 2 ~in. Propellant, J. 

McMorris, Aug. 20, 1942. 13 pp., 2 tables, 5 ill. 

Influence of Additive Agents upon the Burning Characteristics of Extruded Grains of 
Double-Base Propellant, P. A. Longwell, B. H. Sage, and W. N. Lacey, June 15, 
1942. 8 pp., 2 tables, 2 ill. 

Variation in Weight of Grains of Fixed Length for Mousetrap Ammunition, Sept. 3, 
1942. 7 pp., 2 tables, 2 ill. 

Specifications for Propellant of Mousetrap Ammunition, Aug. 19, 1942. 8 pp., 
4 ill. 

Description and Tentative Specifications for Loading the Motor of Mousetrap Ammuni- 
tion, Sept. 8, 1942. 32 pp., 1 table, 14 ill. 

Description and Tentative Specifications for Loading the Motor of AS Projector Charge, 
Feb. 15, 1943. 31 pp., 2 tables, 17 ill. 

Specification for Loading the Motor of Subcaliber Mousetrap Ammunition , Sept. 8, 
1942. 22 pp., 1 table, 9 ill. 

Specifications for Loading the Motor of 4-5-in. BR Ammunition, Sept. 8, 1942. 32 pp., 
1 table, 14 ill. 

Available Propellant Shapes, July 20, 1942. 30 pp., 28 ill. 

Influence of Size of the Axial Perforation Upon the Performance of Radial-Burning 
Grains, D. S. Clark, W. N. Lacey, and B. H. Sage, Sept. 21, 1942. 15 pp., 1 table, 

7 ill. 

Influence of Further Additive Agents Upon the Burning Characteristics of Extruded 
Double-Base Propellant, B. H. Sage, Nov. 17, 1942. 40 pp., 3 tables, 26 ill. 
Experimental Production Facilities at the Eaton Canyon Site, B. H. Sage, Nov. 24, 
1942. 42 pp., 30 ill. 

Extrusion and Burning Characteristics of a Double-Base Propellant Employing Ethyl 
Centralite as Stabilizer, B. H. Sage and others, Nov. 23, 1942. 15 pp., 4 tables, 
6 ill. 

Investigations of Double-Base Powders, R. B. Corey, R. E. Escue, A. L. LeRosen, 
H. A. Levy, and L. Pauling, July 1942. 16 pp. 

Specifications for CWB Propellant Grain, Igniter, and Motor Loading, Nov. 30, 1942. 
10 pp. 

Testing of Quality of Solvent Extruded Powder, B. H. Sage, Feb. 5, 1943. 9 pp., 2 
tables, 4 ill. 


330 


BIBLIOGRAPHY 


cit No. 
IDC 27 

IDC 28 

IDC 29 

IDC 30 

IDC 31 
IDC 32 

IDC 33 

IDC 34 

IDC 35 

IDC 36 

IDC 37 

IDC 37 
Rev. 
IDC 38 

IDC 39 

IDC 40 

IDC 41 

IDC 42 

IDC 43 

IDC 44 

IDC 45 

IEC 1 

IEC 1.2 

IEC 2 

IEC 3 
IEC 4 

IEC 5 

IEC 6 
IEC 7 
IEC 8 

IEC 9 

IEC 11 


NDRC No. OSRD No. 

Corrosion Tests on Various Metals in Solutions of Smokeless Powders , B. H. Sage, 
Feb. 20, 1943. 8 pp., 2 tables. 

Testing of Quality of Extruded Grains Received from Bruceton, B. H. Sage and R. N. 

Wimpress, Apr. 6, 1943. 7 pp., 4 tables, 1 ill. 

Manufacture and Performance of Grains for Use in a 2.5-in. Reaction Chamber , B. H. 

Sage, Apr. 10, 1943. 16 pp., 2 tables, 10 ill. 

Methods of Calculation of Results of Static Proof -Firing Tests at CIT, B. H. Sage. 

5 pp., 1 ill. 

Status of Pilot Plant on May 81, 1943 , B. H. Sage. 11 pp., 5 ill. 

Investigation of the Effectiveness of Flash Protection Afforded by Process Storage Com- 
partments, B. H. Sage, July 10, 1943. 23 pp., 18 ill. 

Ignition of the Powder Charge in an Extrusion Press , B. H. Sage, May 5, 1943. 32 
pp., 14 ill. 

Design of a CWR grain for the 3.25-in. Mk 5 Motor , Q. Elliott and B. H. Sage, Aug. 5 
1943. 17 pp., 8 tables, 7 ill. 

A Study of Certain Hazards Involved in the Loading and Assembly of Rocket Motors, 
A. D. Ayers and B. H. Sage, July 15, 1943. 15 pp., 13 ill. 

Preliminary Studies of the Ballistic Behavior of 218 B Propellant, B. H. Sage, Sept. 
4, 1943. 16 pp., 5 tables, 9 ill. 

Partially Colloided Double-Base Powder Manufactured by a Solvent Process, B. H. 

Sage, Sept. 10, 1943. 20 pp., 4 tables, 13 ill. 

Partially Colloided Double-Base Powder Manufactured by a Solvent Process, B. H. 

Sage, May 23, 1945. 15 pp., 5 tables, 9 ill. 

Changes Occurring in the Propellant Grain After Extrusion ( Temperature Storage 
History), B. H. Sage, Sept. 28, 1943. 11 pp., 4 tables, 4 ill. 

Ignition of Ballistite during Confined Pressing Operations, B. H. Sage, Sept. 20, 1943. 
19 pp., 14 ill. 

A Method of Analysis for Gas-Phase Mixtures of Acetone, Water, and Air, B. H. Sage, 
Mar. 13, 1944. 5 pp., 3 tables, 1 ill. 

2126 Effect of Aluminum on Burning Properties of Solventless Ballistite, B. H. Sage, May 
4, 1944. 14 pp., 8 tables, 13 ill. 

Study of Partial Burnings of Mk 6, Mk 7, and Mk 8 Grains, C. T. Elvey,C.D. Swan- 
son, and C. E. Duemler, July 25, 1944. 26 pp., 3 tables, 26 ill. 

History of Solventless Extrusion of Double-Base Propellant at the California Institute 
of Technology, R. N. Wimpress, Mar. 1, 1945. 24 pp. 

Development of a Hexaform Propellant Grain for 3.25-in. Rocket Motors, B. H. Sage, 
Feb. 1, 1946. 33 pp., 6 tables, 19 ill. 

Photographic Investigation of the Reaction of Rocket Propellant Grains, B. H. Sage, 
Feb. 15, 1946. 20 pp., 13 ill. 

Report as . to Feasibility of Visual Coincidence Scoring by Two Observers of Tracer Bul- 
lets Shot at Rocket Targets, J. W. M. DuMond, June 6, 1942. 5 pp. 

Report as to Utility of Visual Coincidence Scoring by Two Observers of Tracer Bullets 
Shot at Rocket Targets, J. W. M. DuMond, Apr. 29, 1942. 10 pp., 2 ill. 

On the Use of Compressed Gas for Counterpoising Mobile Multiple Rocket Projectors, 
J. W. M. DuMond, Feb. 24, 1942. 10 pp. 

Proof Test of PBY VB-Projector, F. C. Lindvall, Dec. 14, 1942. 9 pp., 6 ill. 

Packing Box Projectors, F. C. Lindvall, F. Fredericks, and P. E. Lloyd, Dec. 17, 

1942. 7 pp., 5 ill. 

The Status of 16-Channel CWB Projector as of Jan. J, 1943, F. C. Lindvall, Jan. 9, 

1943. 7 pp., 5 ill. 

Use of A/S Projector Mark 20, L. B. Slichter, Feb. 17, 1943. 26 pp. 

Performance of the “ Fishtail” Projector, June 1, 1943. 5 pp., 4 tables, 1 ill. 
Standardization Tests of CIT Type 1 Mod 1 4-5-in. BR Launcher: Tests at Goldstone, 

6 April, 8 July, 15 July, 26 July; Test at Pendleton 29 July 1943. 4 pp. 
Standardization Tests of CIT Type 3 Launcher (Wooden 3-Rail for 4-5-in. BR), Aug. 

7, 1943. 8 pp., 2 tables, 3 ill. 

CIT Type 33 Universal Emplacement Launcher, H. A. Meneghelli, Mar. 1, 1944. 
3 pp., 5 ill. 

Manual: Description and Instructions for Use, CIT Type 31C Launcher, H. A. Men- 
eghelli, May 1, 1944. 45 pp., 1 table, 29 ill. 


IEC 12 


NDRC No. OSRD No. 


BIBLIOGRAPHY 


331 


CIT No. 
IEC 13 

IEC 14 
IEC 16 
Prel. 

IFC 1 

IFC 2 

IGC 1.1 
IGC 1.2 
IGC 1.3 
IGC 1.4 

IGC 2 

IGC 3 

IGC 4 

IGC 5 

IGC 6 

IGC 7 
IGC 8 

IGC 10 
IGC 11 

IHC 1.1 

IHC 1.2 

IHC 1.3 

IHC 1.4 

IHC 2 

IHC 3 

IHC 4 
IHC 6 
IHC 7 
IHC 8 
IHC 9 
IHC 10 
IHC 11 
IHC 12 
IHC 13 
IHC 14 
IHC 15 
IHC 16 
IHC 17 

IHC 18 

IHC 19 
IHC 20 


Instructions and Drawings for Installation of Mk 5 Mod 1 Launchers on P-47 Air- 
craft , June 23, 1944. 13 pp., 8 ill. 

Abridged Catalog Entry, CIT Launcher Type 49, Aug. 10, 1944. 15 pp., 12 ill. 
Stresses in Catapults for Launching Projectile Models, L. B. Slichter and J.G. Wendel, 
Sept. 12, 1945. 33 pp., 3 tables, 24 ill. 

The CIT Camera for Measuring Accelerations and Velocities of Projectiles , I. S. Bowen 
and J. B. Edson, Feb. 16, 1942. 15 pp., 1 table. 

Photographic Methods for Determining the Performance of Rockets Fired in a Forward 
Direction from Airplanes, I. S. Bowen, Sept. 15, 1943. 4 pp. 

Static Firing Tests, W. N. Arnquist, Dec. 24, 1941. 11 pp., 5 tables, 1 ill. 

Static Firing Tests, W. N. Arnquist, Dec. 29, 1941. 3 pp., 1 ill. 

Static Firing Tests, W. N. Arnquist, Jan. 19, 1942. 4 pp., 1 ill. 

Static Firing Tests on Extruded Propellant, W. N. Arnquist, Jan. 22, 1942. 7 pp., 
6 ill. 

Investigation of Strain-Sensitive Resistance Pressure Gages and Associated Recording 
Equipment, D. S. Clark and R. W. Alcock, May 13, 1942. 19 pp., 5 tables, 7 ill. 
Performance of Static Firing Equipment as Influenced by the Length of Pressure Line 
from Bourdon Coil to Test Motor, E. L. Ellis, Dec. 5, 1942. 6 pp., 2 tables, 3 ill. 
Rocket MotorTube Bending Machine for Decreasing CG Malalignment, L. A. Richards, 
Mar. 1, 1943. 6 pp., 3 tables, 4 ill. 

Determination of Head Malalignment of Rocket Projectiles, J. G. Waugh, Apr. 10, 
1943. 7 pp., 10 ill. 

Tentative Inspection Procedures for Extruded Ballistite Grains, B. H. Sage, Aug. 25, 
1943. 32 pp., 3 tables, 11 ill. 

A Study of Nozzle Erosion, Motor Test Group, Mar. 8, 1944. 3 pp., 13 ill. 

Theory of the Oil Line with Applications to the Interpretation of Pressure-Time Record- 
ings, B. and D. Locanthi, Jan. 25, 1945. 26 pp., 4 tables, 14 ill. 

An Analysis of the Bourdon Coil, B. Locanthi, June 27, 1945. 8 pp., 5 ill. 

Equipment and Procedures Used in the Static Firing of Rocket Motors at Eaton Canyon, 
B. H. Sage, Mar. 1, 1946. 43 pp., 47 ill. 

Firing of S x /i-in. UP’s at Eaton Canyon Site on Oct. 15, 1941, W. A. Fowler, Oct. 
28, 1941. 7 pp., 3 tables, 1 ill. 

Firing of 6-in. Antisubmarine UP’s at Eaton Canyon Site on Oct. 19 and 26, 1941. 
W. A. Fowler, 11 pp., 2 tables, 5 ill. 

Interpolations of Range Measurements with the Mule, W. N. Arnquist, Feb. 5, 1942. 
10 pp., 2 tables, 5 ill. 

Tests on Nozzle Performance with the Mule, W. N. Arnquist, Mar. 11, 1942. 4 pp., 
1 ill. 

Impact Decelerometers, R. L. Noland, T. H. Wianko, W. P. Huntley, and B. H. 
Rule, Nov. 13, 1943. 36 pp., 19 ill. 

Proposed Layout of Facilities for Naval Ordnance Test Station, Inyokern, California, 
Nov. 12, 1943. 22 pp., 1 ill. 

Measurement of Deceleration Components, R. Stokes, Dec. 20, 1943. 2 pp. 

The Coil Spring Decelerometer , R. Stokes, Jan. 3, 1944. 2 pp., 1 ill. 

The Navy Type Spring Decelerometer, R. Stokes, Jan. 3, 1944. 4 pp., 1 table, 3 ill. 
The Multi-Unit Indenter Decelerometer, R. Stokes, Jan. 3, 1944. 3 pp., 2 ill. 

DeForest Scratch Gage Decelerometer, R. Stokes, Jan. 3, 1944. 2 pp., 1 ill. 

Mercury Decelerometer, R. Stokes, Jan. 3, 1944. 5 pp., 4 ill. 

The Double Diaphragn Decelerometer , R. Stokes, Dec. 20, 1943. 5 pp., 4 ill. 

The Copper Ball Decelerometer , R. Stokes, Jan. 3, 1944. 4 pp., 3 ill. 

The Rotary Disk Decelerometer, R. Stokes, Jan. 3, 1944. 2 pp., 1 ill. 

Recording Step Decelerometer, R. Stokes, Dec. 20, 1943. 6 pp., 4 ill. 

Projectile Dolly for Carrier Use, G. M. Mosteller, Mar. 14, 1944. 44 pp., 42 ill. 

1-in. Model Test Facilities, B. H. Rule and R. C. Bradley, May 6, 1944. 3 pp., 9 ill. 
Modeling of Water Entry of Bombs and Projectiles, L. B. Slichter, Mar. 31, 1944. 
30 pp., 1 table, 1 ill. 

Sound Ranging Facilities and Procedures at the CIT Torpedo Launching Range, F. R. 

Watson, May 15, 1944. 17 pp., 19 ill. 

Status of the China Lake Pilot Plant, about 50 pp. 

The Aircraft Rocket Training Aid, Jan. 15, 1945. 5 pp., 4 ill. 


332 


BIBLIOGRAPHY 


CIT No 
IHC 21 
Prel. 
IIC 1 

IIC 2 
IIC 3 

IIC 4 

IIC 5 

IIC 6 

IIC 7 
IIC 8 

IIC 9 

IIC 10 

IIC II 

IIC 14 
IIC 15 
IIC 16 
IIC 17 

IIC 18 

IIC 20 

IIC 21 

IKC 1 
ILC 1 

ILC 2 

ILC 3 
ILC 4 

ILC 5 

IMC 1 

IMC 2 

IMC 3 

INC 1 

INC 2 
INC 3 

INC 4 

INC 5 

INC 6 


NDRC No. OSRD No. 

Underwater Photography Facilities at Morris Dam , B. H. Rule, Mar. 6, 1945. 5 pp. 

Characteristics of HIR Fuze for Mousetrap Ammunition , T. Lauritsen, June 3, 1942. 
8 pp., 2 ill. 

Propeller Arming Fuze , N. Gunderson, Aug. 31, 1942. 10 pp., 1 table, 4 ill. 

Tests of Type HIR Fuzes, B. H. Rule and W. P. Huntley, Sept. 14 to Nov. 9, 1942. 
14 pp., 12 tables, 2 ill. 

Tests of Type HIR Fuze Modified to Increase Firing Sensitivity, B. H. Rule and 
W. P. Huntley, Oct. 12 to Dec. 1, 1942. 13 pp., 19 tables, 5 ill. 

Firing Tests of ABN-7 A Magnetic Fuze, B. H. Rule and W. P. Huntley, Dec. 31, 
1942. 4 pp., 1 ill. 

Arming and Firing of Mark No. 31 Fuze, B. H. Rule and W. P. Huntley, Dec. 30, 

1942. 5 pp., 1 table, 1 ill. 

IHR Fuze Tests, B. H. Rule and W. P. Huntley, Dec. 31, 1942. 3 pp. 

Tests of Preliminary Firing Mechanism for SIR Fuze, B. H. Rule and W. P. Huntley, 
Jan. 1943. 3 pp., 1 table. 

SIR Fuze Tests, Dec. 8, 1942 to Jan. 15, 1943, N. Gunderson, D. E. Brink, C. F. 

Robinson, V. Rasmussen, and R. B. King. 3 pp. 

Preliminary Tests of Mark 24 Mine and Proposed Fuzes, M. Mason, Feb. 1943. 4 pp., 
2 tables. 

Tests of Production Fuzes of Mark 31 Mod 1 Fuzes, B. H. Rule and W. P. Huntley, 
Feb. 1943. 7 pp., 3 tables. 

Tests of SIR {Mark 129) Fuze. Feb. 27 to April 27, 1943. 7 pp., 1 ill. 

Mark 138 Fuze Tests, B. H. Rule and W. P. Huntley, June 8, 1943. 2 pp., 1 table. 
Horn Fuze Tests, B. H. Rule and W. P. Huntley, July 8, 1943. 6 pp., 4 tables. 

Soft Metal Wires in Shear as Delay Elements for Time Fuzes, J. McMorris, July 12, 

1943. 11 pp., 1 table, 7 ill. 

The Mark 140 Fuze {HIR 3): Tests of Arming Depth, Premature Firing, and Sensi- 
tivity, Fuze Group. 4 pp. 

XMTF Fuze {Experimental, BuOrd Mousetrap ) Morris Dam Report No. 89, B. H. 

Rule and W. P. Huntley, Oct. 27, 1943. 9 pp., 3 tables, 2 ill. 

Mk I 46 Fuze {PIR): Static Firing Progress Report, O. E. Brink (Fuze Group), 
Mar. 18, 1944. 3 pp., 9 ill. 

Smoke Tracers for UP’s, J. McMorris, Feb. 14, 1942. 3 pp. 

Study of Nozzle Side Forces by Means of Compressed Air Jet, G. E. Kron and O. C. 

Wilson, Dec. 15, 1942. 9 pp., 3 tables, 4 ill. 

Further Investigations Conducted with the Yaw Machine, Accuracy Committee, 
Jan. 4, 1943. 3 pp., 1 table, 1 ill. 

The Computation of Pressure-Time Curves, C. T. Elvey, June 3, 1943. 21 pp., 16 ill. 
Internal Ballistic Calculations for Four Motors with Neutral-Burning Grains, C. T. 
Elvey, June 8, 1943. 5 pp., 4 ill. 

Temperature Gradient in the Tubing of the 3.25-in. Rocket Motor Mk 7, F. E. Roach, 
J. M. Schmidt, and W. F. Nash, Jr., May 22, 1944. 5 pp., 2 tables, 4 ill. 

Hazards in the Shipment of Propulsive and Nonpropulsive Rockets, June 22, 1943. 
7 pp., 4 tables, 1 ill. 

Developments, Projects and Special Problems, Past, Present and Future {Contract 
OEMsr-418 ), June 1943. 17 pp. 

Estimation of Lot Quality in Sampling by Attributes, J. G. Waugh, Feb. 21, 1944. 
1 p., 3 ill. 

Scoring Register and Recorder for Target Practice, J. Edson, Apr. 13, 1942. 8 pp., 
6 ill. 

Rocket Torpedo Projects, J. Edson, Oct. 27, 1942. 7 pp. 

Examination of Rules for Lead Angles as Affecting the Efficiency of Mousetrap Attacks, 
M. A. Biot, Apr. 8, 1943. 14 pp., 2 tables, 7 ill. 

Dispersed Target Method of Determining Barrage Probabilities, C. G. Willis and 
G. Stromberg, Apr. 8, 1944. 16 pp., 16 ill. 

The Effective Rocket Temperature Indicator, R. White and J. Schmidt, Dec. 12, 1944. 
10 pp., 9 ill. 

An Approximate Method of Computing Hit Probabilities for Rectangular Barrage 
Patterns, N. A. Haskell, Aug. 24, 1945. 11 pp., 6 ill. 


NDRC No. OSRD No. 


BIBLIOGRAPHY 


333 


CIT No. 
IOC 1 

IOC 2 
IOC 3 
IOC 4 
IOC 5 
IOC 6 

IOC 7 

IOC 9 
IOC 10 
IOC 11 
IOC 12 
IOC 13.1 
IOC 13.2 
IOC 13.3 
IOC 13.4 
IOC 13.5 
IOC 13.6 
IOC 13.7 
IOC 14 
IOC 15 
IOC 16 
IOC 17 
IOC 18 
IOC 19 
IOC 20 

IOC 20 
1st rev. 

IOC 21 


Underwater Tests of Mark 6 and Mark 9 Depth Charges, and Depth Charge Pistols 
Mark 6, Mark 6 Mod 1, and Mark 7, B. H. Rule and W. P. Huntley, Mar. 12, 
1943. 27 pp., 11 tables, 20 ill. 

Firing Depth Test of Modified Mark VI Mod 1 Depth Charge Pistols in Mark IX 
Depth Charges, B. H. Rule and W. P. Huntley, May 18, 1943. 4 pp., 2 tables. 
Underwater Performance Tests of Mark IX Depth Charges with Baranol Filling, 
B. H. Rule and W. P. Huntley, May 27, 1943. 8 pp., 2 tables, 3 ill. 

Underwater Tests of Proposed Practice Hedgehog Ammunition, B. H. Rule and W. P. 
Huntley, June 2, 1943. 3 pp., 1 table. 

Slat Deck Impact Deceleration Tests, ASPC Mk 1, B. H. Rule and W. P. Huntley, 
July 12, 1943. 4 pp., 1 table, 2 ill. 

Underwater Performance Tests of Bureau of Ordnance 5-in./38 Caliber Projectile, 
B. H. Rule and W. P. Huntley, Aug. 5, 1943. Morris Dam Report No. 72. 7 pp., 

1 table, 3 ill. 

Underwater Performance Tests of Mark 17 Aircraft Depth Charge with and without 
Nose Projection, B. H. Rule and W. P. Huntley, Aug. 5, 1943. Morris Dam 
Report No. 78. 3 pp., 1 ill. 

Underwater Performance Tests of Bureau of Ordnance Mk 6 Hedgehog Projectile, Aug. 

25, 1943. Morris Dam Report No. 81. 4 pp., 1 table, 2 ill. 

Mark 24 Mine Target Tests, B. H. Rule and W. P. Huntley, Sept. 4, 1943. Morris 
Dam Report No. 65. 8 pp., 1 table, 5 ill. 

BuOrd Subcaliber Ammunition for A S Projector Mark 10 ( Hedgehog ) Underwater 
Performance Tests, B. H. Rule and W. P. Huntley, Sept. 21, 1943. 4 pp., 2 tables. 
^ 2 -Scale Model British Projectile Type C, B. H. Rule and W. P. Huntley, Nov. 1, 
1943. Morris Dam Report No. 90. 15 pp., 2 tables, 11 ill. 

Preliminary Tests of 8-in. -Diameter Model Torpedoes, Memo 1, B. H. Rule and 
W. P. Huntley, Oct. 30, 1943. Morris Dam Test No. 84. 7 pp., 3 tables, 1 ill. 
Preliminary Tests of 8-in. -Diameter Model Torpedoes, Memo 2, B. H. Rule and W. P. 

Huntley, Oct. 30, 1943. 8 pp., 3 tables, 1 ill. 

Preliminary Tests of 8-in. -Diameter Model Torpedoes, Memo 3, B. H. Rule and W. P. 

Huntley, Dec. 1, 1943. 12 pp., 5 tables, 4 ill. 

Preliminary Tests of 8-in. -Diameter Model Torpedoes, Memo J, B. H. Rule and W. P. 

Huntley, Dec. 3, 1943. 15 pp., 5 tables, 8 ill. 

Preliminary Tests of 8-in -Diameter M odel Torpedoes, Memo 5, B. H. Rule and W. P. 

Huntley, Jan. 28, 1944. 14 pp., 4 tables, 5 ill. 

Preliminary Tests of 8-in. -Diameter Model Torpedoes, Memo 6, B. H. Rule and W. P. 

Huntley, Mar. 7, 1944. 13 pp., 4 tables, 4 ill. 

Preliminary Tests of 8-in. -Diameter Model Torpedoes, Memo 7, B. H. Rule and W. P. 

Huntley, Sept. 14, 1944. 21 pp., 2 tables, 11 ill. 

Observations on the Water Entry of a Torpedo, R. W. Ager, Nov. 23, 1943. 14 pp., 
9 ill. 

2 l / 2 -in .-Diameter Model Mk 13 Aircraft Torpedo Tests, B. H. Rule and W. P. Hunt- 
ley, Nov. 15, 1943. Morris Dam Report No. 74. 10 pp., 8 ill. 

Preliminary Tests of l-in.-Diameier Torpedo Models, B. H. Rule and W. P. Huntley, 
Nov. 26, 1943. Morris Dam Report No. 87. 16 pp., 1 table, 11 ill. 

Underwater Tests of Mousetrap with Line and Drogue, B. H. Rule and W. P. Huntley, 
Aug. 12, 1943. Morris Dam Report No. 73. 15 pp., 11 ill. 

Responses of the Control Surfaces of the Mk 13-2 Torpedo, Tests, Calculations, and 
Observations, R. W. Ager, Nov. 15, 1943. 6 pp., 2 tables, 20 ill. 

Mark 12 Depth Charge, B. H. Rule and W. P. Huntley, Nov. 24, 1943. Morris Dam 
Report No. 92. 7 pp., 1 table, 3 ill. 

2282 Water Entry of 8-in. -Diameter Model Aircraft Torpedoes with Hemispherical and 
Ogival Noses, B. H. Rule and W. P. Huntley, Dec. 31, 1943. Morris Dam Report 
No. 93. 54 pp., 12 tables, 26 ill. 

Water Entry of 8-in. -Diameter Aircraft Torpedo Models with Henispherical and 
Ogival Nose, B. H. Rule and W. P. Huntley, Nov. 13, 1944. 42 pp., 12 tables, 
27 ill. 

Underwater Performance of Mk 12 Depth Charge, B. H. Rule, Jan. 26, 1944. 7 pp., 

2 tables, 1 ill. 


334 


BIBLIOGRAPHY 


CIT No. 
IOC 22 


IOC 23 


IOC 24 

IOC 24 
1st rev. 

IOC 25 

IOC 26 


IOC 26.2 


IOC 27 


IOC 28 

IOC 28.2 
Prel. 

IOC 28.2 

IOC 28.3 

IOC 28.4 
Prel. 

IOC 28.4 

IOC 28.5 
Prel. 

IOC 28.5 

IOC 28.5 
Rev. 1 
IOC 29 

IOC 30x 
IOC 30. 2x 
Prel. 

IOC 31 

IOC 31 
Revised 
Prel. 

IOC 31 
Rev. 

IOC 32 
Prel. 

IOC 32 


IOC 33 
Prel, 


NDRC No. OSRD No. 

Underwater Performance of Mk 12 Depth Charge with Mk lJfi Fuze with and without 
Protective Cap, R. L. Noland, Jan. 31, 1944. Morris Dam Report No. 96. 15 pp., 
4 tables, 6 ill. 

Underwater Performance Tests of BuOrd Mk 6 and BuOrd Mk 8 Projector Charges 
{Hedgehog) with Mk 1 40 Fuze and Protective Cap. Morris Dam Report No. 100, 
R. L. Noland and B. H. Rule, Feb. 28, 1944. 2 pp., 1 table, 4 ill. 

2255 Water Entry of 8-in. -Diameter Model Aircraft Torpedoes with Special Noses and Ring 
Tails, B. H. Rule and W. P. Huntley, Apr. 7, 1944. 50 pp., 11 tables, 31 ill. 

Water Entry of 8-in -Diameter Model of Aircraft Torpedoes with Special Noses and 
Ring Tail, B. H. Rule and W. P. Huntley, Nov. 1, 1944. 43 pp., 11 tables, 
29 ill. 

Underwater Performance of Hedgehog with Line and Weight (“ Mustard Plaster”), 
R. L. Noland, May 5, 1944. 16 pp., 10 ill. 

Underwater Performance of 7 . 2-in. -Diameter Fast-Sinking Depth Charge with Mark 
140 Fuze with and without Protective Cap. Morris Dam Report No. 97, B. H. Rule 
and W. P. Huntley, May 13, 1944. 3 pp., 2 tables, 7 ill. 

2223 Underwater Performance of the 7 .2-in. -Diameter Fast-Sinking Depth Charge with Case 
Length Increased 1 in. and 3 in. with Mk 140 Fuze with and without Protective Cap, 
B. H. Rule and W. P. Huntley, Sept. 7, 1944. 8 pp., 1 table, 5 ill. 

Underwater Performance of 6-in. -Diameter Mark 12 Fast-Sinking Depth Charge with 
Tails of Various Sizes and with Mark 140 Fuze and Protective Cap, B. H. Rule 
and W. P. Huntley, June 7, 1944. 14 pp., 2 tables, 10 ill. 

British Type U C” Projectile, B. H. Rule and W. P. Huntley, June 20, 1944. 21 pp., 
4 tables, 10 ill. 

British Type “C” Production Units, B. H. Rule and W. P. Huntley, Sept. 20, 1944. 
14 pp., 2 tables, 11 ill. 

2270 British Type “C” Production Units, B. H. Rule and W. P. Huntley, Sept. 20, 1944. 
12 pp., 2 tables, 11 ill. 

Sinking Velocity and Forward Travel of British Type “C” Projectile, Max Mason 
and L. B. Slichter, Oct. 12, 1944. 

British Type “C” Production Units, Max Mason and L. B. Slichter, Nov. 30, 1944. 
23 pp., 4 tables, 8ill. 

2339 Underwater Trajectory Tests of Production Units of British Projectile Type “C,” 
Max Mason and L. B. Slichter, Jan. 2, 1945. 18 pp., 4 tables, 8 ill. 

Underwater Performance of British Projectile Type “C,” Max Mason and L. B. 
Slichter, May 9, 1945. 36 pp., 11 tables, 13 ill. 

2475 Underwater Performance of British Projectile Type “C,” Max Mason and L. B. 
Slichter, Feb. 12, 1945. 26 pp., 3 tables, 10 ill. 

Underwater Performance of British Projectile Type “C,” Max Mason and L. B. 
Slichter, May 10, 1945. 33 pp., 5 tables, 13 ill. 

Instruction Panphlet: Shroud Ring, Mk 1 Mod 0 as Applied to the Mk 13-Type Tor- 
pedo, June 26, 1944. 12 pp., 8 ill. 

Large Bomb for B-17 Plane {Project “Egg”), July 29, 1944. 14 pp., 10 ill. 

Project EGO — Nose Impact Tests, Max Mason and L. B. Slichter, Jan. 22, 1945. 
14 pp., 1 table, 27 ill. 

2374 The Effect of Roughness of Sea on the Entry Angle of a Projectile, R. I. Piper, Aug. 19, 
1944. 5 pp. 

The Effect of Roughness of the Sea on the Entry and Ricochet Angles of a Projectile, 
Max Mason and L. B. Slichter, Jan. 30, 1945. 6 pp., 1 table. 

The Effect of Roughness of the Sea on the Entry and Ricochet Angles of a Projectile, 
Max Mason and L. B. Slichter, Feb. 1, 1945. 6 pp., 1 table, 3 ill. 

1-in. Model Results Concerning the Effect of Pitch and Entry Angle on the Underwater 
Orbit of Mark 13-2 Torpedo, M. Mason and L. B. Slichter, Aug. 19, 1944. 18 pp., 
2 tables, 5 ill. 

2198 The Effect of Pitch and Entry Angle on the Underwater Orbit of Mark 13-2 Torpedo 
as Determined by Tests with 1.0-in. Model, M. Mason and L. B. Slichter, Aug. 28 
1944. 12 pp., 2 tables, 7 ill. 

Bending Stresses in a Bomb Case at Water Impact, N. A. Haskell, Aug. 25, 1944. 9 pp 


BIBLIOGRAPHY 


335 


C1T No. 
IOC 34 
Prel. 
IOC 34 

IOC 35 

IOC 36 

IOC 37 


IOC 37 
1st rev. 

IOC 38 
Prel. 

IOC 38 
1st rev. 

IOC 39 
Prel. 

IOC 39 
1st rev. 
IOC 40 
Prel. 

IOC 40.2 
Prel. 

IOC 41 

IOC 42x 
Prel. 

IOC 43 
Prel. 

IOC 44 
Prel. 

IPC 1 

IPC 2 

IPC 3 

IPC 4 

IPC 5 


IPC 6 

IPC 6 
Addendum 1 
IPC 7 

IPC 8 

IPC 9 

IPC 10 


NDRC No. OSRD No. 

Air Flight Tests of 8-in. -Diameter Model Torpedoes , B. H. Rule and W. P. Huntley, 
Sept. 14, 1944. 17 pp., 2 tables, 12 ill. 

2227 Air Flight Tests of 8-in. -Diameter Model Torpedoes, B. H. Rule and W. P. Huntley, 
Sept. 26, 1944. 15 pp., 2 tables, 12 ill. 

Use of the Hydropressure Plug for Water Impact Studies, R. R. Stokes, Sept. 15, 1944. 
8 pp., 8 ill. 

Tests of Heavy Pod Dummy Rugged Model, B. H. Rule and W. P. Huntley, Oct. 4, 
1944. 19 pp., 6 tables, 11 ill. 

Correlation Between 1 -in. -Diameter Model and Prototype Mark 13-2 Dummy Torpedo, 
F. C. Lindvall, Max Mason, and L. B. Slichter, Nov. 17, 1944. 14 pp., 2 tables, 
5 ill. 

2320 Correlation Between 1 -in. -Diameter Model of Mark 13-2 Dummy Torpedo and the 
Prototype, F. C. Lindvall, Max Mason, and L. B. Slichter, Dec. 15, 1944. 11 pp., 
2 tables, 5 ill. 

A Test of Modeling of the Underwater Trajectory of Cone-Nosed Cylinders of %-in.- 
to 5-in. -Diameter, Max Mason and L. B. Slichter, Mar. 24, 1945. 26 pp., 2 tables, 
21 ill. 

2468 A Test of Modeling of the Underwater Trajectory of Cone-Nosed Cylinders Y%- in- 
to 5 -in. -Diameter, Max Mason and L. B. Slichter, Apr. 2, 1945. 23 pp., 2 tables, 
20 ill. 

Effect of Nose Shape upon Underwater Trajectory of Mark 13-2 Torpedo Models, 
Max Mason and L. B. Slichter, Nov. 22, 1944. 32 pp., 41 ill. 

2326 Effect of Nose Shape on Underwater Trajectory of Mk 13-2 Torpedo Models, Max 
Mason and L. B. Slichter, Dec. 22, 1944. 28 pp., 43 ill. 

Water Entry Tests of NAE Beacon Mark 1 Mod 1, Max Mason and L. B. Slichter, 
Mar. 1, 1945. 6 pp., 1 table, 1 ill. 

Water Entry Tests of NAE Beacon Mark 1, Mods 1 and 2, Max Mason and L. B. 
Slichter, Aug. 14, 1945. 2 pp. 

2467 Determination of Depth of Dive of Aircraft Torpedoes with the Foxboro Depth and Roll 
Recorder Mk 1 Mod 2, F. C. Lindvall, May 28, 1945. 21 pp., 4 tables, 14 ill. 

Underwater Trajectory and Pitch Sensitivity of 1 -in. -Diameter Model IJP Torpedo, 
Max Mason and L. B. Slichter, Aug. 18, 1945. 3 pp., 3 tables, 10 ill. 

Protective Cage for Afterbody of Mk 13-6 Torpedo, Max Mason and L. B. Slichter, 
Aug. 27, 1945. 15 pp., 3 tables, 8 ill. 

Underwater Performance of 3-in. -Diameter Mk 15 Depth Charge, Max Mason and 
L. B. Slichter, Sept. 21, 1945. 10 pp., 3 tables, 3 ill. 

UP Vacuum Ranges as Functions of the Burning Distance, W. N. Arnquist, May 31, 
1942. 3 pp. 

Effect of Tail Spin on Underwater Projectile, Dec. 9, 1941 to Jan. 10, 1942, B. H. 
Rule. 6 pp., 1 table, 1 ill. 

Performance Characteristics of 7 -in -Diameter Underwater Bombs at Velocities up to 
50 ft/sec Mar. 2 to May 15, 1942, B. H. Rule. 28 pp., 17 ill. 

Performance Characteristics of 8-in. -Diameter Underwater Bombs at Velocities up to 
50 ft/sec, Feb. 24 to June 15, 1942, B. H. Rule. 49 pp., 2 tables, 32 ill. 

Underwater Characteristics of 7 .2-in. -Diameter Mousetrap Projectile and 7 -in. -Diam- 
eter Integral Motor Modification Tests Mar. 31 to May 15, 1942 , B. H. Rule. 
49 pp., 4 tables, 30 ill. 

Performance Characteristics of 2 Y-in. -Diameter Underwater Integral Motor Target 
Bomb, B. H. Rule, May 29, 1942. 12 pp., 4 tables, 6 ill. 

Performance Tests of Modifications of 2 Y-in. -Diameter Subcaliber Target Bomb with 
Ring Tail Assembly, B. H. Rule, Sept. 2, 1942. 9 pp., 3 tables, 5 ill. 

Performance Comparisons for 7 .2-in -Diameter Modified Mousetrap Projectile, May 
28 to June 20, 1942, B. H. Rule, 23 pp., 4 tables, 10 ill. 

Model Tests of Mark VI Depth Charge with Nose and Tail Adapters June 8 to 20, 1942, 
B. H. Rule. 15 pp., 1 table, 9 ill. 

Effect of Tail Spin on Underwater Projectiles, B. H. Rule, Aug. 5, 1942. 24 pp., 1 
table, 11 ill. 

The Effect of Fin Size, Burning Time, and Projector Length on the Accuracy of UP’s, 
I. S. Bowen and L. Blitzer, Nov. 13, 1942. 9 pp., 1 table, 3 ill. 



336 


BIBLIOGRAPHY 


CIT No. 
IPC 11 

IPC 13 

IPC 14 

IPC 15 

IPC 16 

IPC 18 

IPC 19 

IPC 20 
IPC 21 

IPC 22 

IPC 23 

IPC 24 
IPC 25 

IPC 26 
IPC 27 

IPC 27.1 

IPC 28 

IPC 29 

IPC 30 

IPC 31 

IPC 32 

IPC 32.1 

IPC 32.2 

IPC 33 

IPC 34 

IPC 35 

IPC 36 

IPC 37 

IPC 37 
Addendum 1 
IPC 38 

IPC 39 


NDRC No. OSRD No. 

Effect of Wind on Mean Deflections of ASB, 4.5-in. BR, and CWB, L. Blitzer, Jan. 
23, 1943. 4 pp.,3 ill. 

Air Flight and Underwater Characteristic Tests of Proposed VAB-7V Vertical Air 
Bombs, B. H. Rule and W. P. Huntley, Jan. 1943. 7 pp., 3 tables, 3 ill. 

Water Entry and Underwater Trajectory Tests on Bureau of Ordnance ASPC ( Mouse- 
trap ), B. H. Rule, Dec. 3, 1942. 5 pp., 1 table, 2 ill. 

Firing Tests of the HHR Fuze, Sept. 18 to Nov. 25, 1942, B. H. Rule and W. P. Hunt- 
ley. 5 pp., 4 tables. 

AIR Fuze, Tested Sept. 9 to Oct. 29, 1942, B. H. Rule and W. P. Huntley. 4 pp., 

1 table. 

Summary of Underwater Performance Data for ASPC Mk 1 , M. Mason, Feb. 1943. 
3 pp., 1 table. 

Procedure in Calculating “Dispersions” and “Probable Errors,” W. A. Fowler, Mar. 
1943. 1 p. 

Procedure in Calculating Gas Malalignments, W. A. Fowler, Mar. 19, 1943. 5 pp. 
Determination of the Vertical Deviations and Relations between the Various Deviations, 
L. Davis, Jr., Mar. 18, 1943. 9 pp., 1 111. 

Effect on Accuracy of Fire of Rotation of a Rocket about its Axis of Symmetry, Ac- 
curacy Committee, Mar. 15, 1943. 6 pp. 

Deceleration Coefficient of the 2-in. A A at High Velocities, L. Davis, Jr., Apr. 22, 
1943. 9 pp., 3 ill. 

Tests of Jiggle Switch Fuze in Mk 24 Mine, Apr. 24, 1943. 4 pp. 

Addition of a Rocket to a Shell to Increase the Velocity, L. Davis, Jr., June 5, 1943. 

7 pp., 4 ill. 

Curves for the Prediction of Ranges, F. E. Roach, June 1, 1943. 11 pp., 10 ill. 

VAR Subcaliber Underwater Performance Tests, B. H. Rule, W. P. Huntley, June 
17, 1943. 6 pp., 2 tables, 3 ill. 

VAR Subcaliber Underwater Performance Tests, B. H. Rule and W. P. Huntley, 
July 15, 1943. 3 pp., 2 tables. 

VAR-7V11 Antisubmarine Rocket, Entry and Underwater Tests, B. H. Rule and 
W. P. Huntley, June 2, 1943. 5 pp., 6 tables. 

Mark IX Mousetrap {Proposed New ASPC), Entry and Underwater Tests, B. H. Rule 
and W. P. Huntley, June 5, 1943. 6 pp., 5 tables. 

VAR-7V13 Vertical Antisubmarine Rocket, Entry and Underwater Test, B. H. Rule 
and W. P. Huntley, June 2, 1943. 5 pp., 6 tables. 

Deflection of Rockets by Wind, W. N. Arnquist and R. J. Kennedy, June 30, 1943. 

2 pp. 

ASPC Subcaliber Underwater Performance Tests, B. H. Rule and W. P. Huntley, 
June 23, 1943. 4 pp., 2 tables. 

ASPC Subcaliber Ammunition Underwater Performance Tests, B. H. Rule and W. P. 
Huntley, July 7, 1943. 2 pp., 1 table. 

ASPC Subcaliber Ammunition Underwater Performance Tests, B. H. Rule and W. P. 

Huntley, Oct. 21, 1943. 6 pp., 1 table, 2 ill. 

Underwater Performance of B. 0. ASPC Mark 11, B. H. Rule and W. P. Huntley, 
Aug. 10, 1943. 13 pp., 3 tables, 7 ill. 

Forward Firing of the British UP 3 and CIT 3 A Rockets from Airplanes, L. Blitzer, 
July 24, 1943. 12 pp., 1 table, 7 ill. 

Method of Changing a Standardization Report when the Weight But Not the Shape of 
a Rocket is Changed, L. Davis, Jr., July 28, 1943. 4 pp. 

Air Drag Test of 2-in. AAR, F. E. Roach and L. Davis, Jr., June 19, 1943. 9 pp., 

8 tables. 

Underwater Test of 2 l / 2 ~in. ASR Subcaliber with Magnesium Flare Head , B. H. Rule 
and W. P. Huntley, Aug. 3, 1943. 4 pp., 1 table, 1 ill. 

Underwater Tests of 2}^-in. ASR Subcaliber with Magnesium Flare Head, B. H. Rule 
and W. P. Huntley, Dec. 27, 1943. 1 p. 

AS Projector Charges — Terminal Velocity Summary, B. H. Rule and W. P. Huntley, 
Aug. 18, 1943. Morris Dam Report No. 83. 2 pp., 1 ill. 

Dispersion of Rockets in Forward Firing from Airplanes with Reference to the CIT 
3A12 and British UP 3 Rockets, L. Blitzer, Aug. 25, 1943. 4 pp. 

a? 


NDRC No. OSRD No. 


BIBLIOGRAPHY 


337 


CIT No. 
IPC 39.2 

IPC 40 

IPC 41 

IPC 42 

IPC 43 

IPC 44 

IPC 45 

IPC 46 

IPC 47 

IPC 48 
IPC 50 
IPC 51 
IPC 52 
IPC 53 
IPC 54 
IPC 55 
IPC 57 

IPC 58 

IPC 59 

IPC 60 

IPC 61 

IPC 62 

IPC 63 

IPC 64 

IPC 65 

IPC 66 

IPC 67 
Prel. 

IPC 67 


Ammunition Dispersion of Long-Burning Unrotated Rockets in Forward Firing from 
Airplanes , L. Blitzer, Apr. 10, 1944 (Revision of IPC 39). 4 pp., 1 ill. 

BuOrd Subcaliber ASPC Underwater Performance Tests, B. H. Rule and W. P. 
Huntley. 5 pp., 3 tables. 

BuOrd Mk 2 Subcaliber VAR Underwater Performance Tests, B. H. Rule and W. P. 
Huntley, Sept. 18, 1943. 5 pp., 3 tables. 

Range Corrections for Altitude and Wind, L. Davis, Jr., and S. Rubin, Oct. 9, 1943. 
5 pp., 1 ill. 

Estimation of the Inherent Dispersion of Rockets by Firing Pairs of Rockets Simul- 
taneously, J. G. Waugh, Oct. 11, 1943. 7 pp., 1 ill. 

Universal Sighting for Guns and Launchers, L. Davis, Jr., Oct. 11, 1943. 2 pp., 3 
tables. 

Approximate Formulae for Calculation of Deflection of Rockets, J. G. Waugh, Nov. 
3, 1943. 2 pp., 1 ill. 

Analysis of Forward Firing Data from TBM and PV-1, L. Blitzer, Sept. 28-29, 

1943. 5 pp., 3 ill. 

Preliminary Tests of 1 -in. -Diameter Model Antisubmarine Bombs and Projectiles 
with Bubble Eliminating Devices, B. H. Rule and W. P. Huntley, Morris Dam 
Report No. 69. 21 pp., 17 ill. 

Theoretical Trajectories for a Projectile Resembling the 5-in. Rocket, L. Davis, Jr., 
Nov. 1943. 2 pp., 1 ill. 

Spinner Theory: Qualitative Notes, Part I, L. Davis, Jr., Dec. 18, 1943. 10 pp., 

2 ill. 

Underwater Tests of 2%-in. VAR Subcaliber with Mg Flare Heads, B. H. Rule and 
W. P. Huntley, Dec. 20, 1943. 3 pp., 1 table. 

Drag Characteristics of Various Aircraft Rocket Projectiles, Hsue-Shen Tsien and 
L. Davis, Jr., Jan. 15, 1944. 5 pp., 4 tables, 2 ill. 

Effect of Mallaunching on the Accuracy of Rockets (Spinners), L. I. Epstein, Jan. 13, 

1944. 3 pp., 2 tables, 1 ill. 

Radius of Curvature of the Underwater Trajectory of a Rocket, L. Davis, Jr., Mar. 17, 
1944. 5 pp., 1 table. 

Formulas for the Spin Produced by Inclined Jets, L. Davis, Jr., Mar. 25, 1944. 
9 pp., 4 ill. 

Estimation of Drag and Nose-Lift Coefficients of Some Rockets Necessary to Give a 
Prescribed Underwater Behavior, L. Davis, Jr., and L. I. Epstein, Apr. 10, 1944. 

3 pp., 2 tables. 

The Choice of a Launcher Angle to Give an Impact Pattern Which Represents the Di- 
rections of Motion at the End of Burning, L. Davis, Jr., May 5, 1944. 3 pp., 2 ill. 
The Theoretical Determination of the Effective Rocket Temperature for the 3.25-in. Mk 

7 Motor When Newton’s Law of Cooling Holds, L. Davis, Jr., May 22, 1944. 6 pp. 
Test of the Significance of Differences in Mean Deviation in Deflection and Mean Devia- 
tion in Range of Rockets, J. G. Waugh, May 25, 1944. 13 pp. 

Experimental Values for Rocket Sight Calibration for Several Aircraft, O. D. Framp- 
ton, June 20, 1944. 10 pp., 4 tables. 

Preliminary: Sighting Data for British Aircraft Rocket, 3.25-in. Motor with 11-lb 
Cruciform Grain, 62-lb Body; also Tubular Grain, June 23, 1944. 9 pp., 9 tables. 
Effect of Tip-Off on the Forward Firing Trajectory of a Fin-Stabilized Rocket, L. I. 
Epstein, June 28, 1944. 9 pp., 1 ill. 

3.0-in. AR Bodies with Nonricochet Properties at Low Angles of Water Impact, I. S. 

Bowen, Oct. 10, 1944. 4 pp., 1 table, 2 ill. 

Underwater Behavior of the 11.75-in. Aircraft Rocket, I. S. Bowen, Oct. 25, 1944. 

8 pp., 1 table, 6 ill. 

Underwater Tests of l^-in. Model of 3.5-in. AR, Max Mason and L. B. Slichter, 
Nov. 6, 1944. 27 pp., 1 table, 26 ill. 

Curvatures and Other Geometrical Characteristics of Selected Nose Types, Max Mason 
and L. B. Slichter, Nov. 25, 1944. 24 pp., 16 ill. 

2350 Curvatures and Other Geometrical Characteristics of Selected Nose Types for Aircraft 
Torpedoes, Max Mason and L. B. Slichter, Jan. 15, 1945. 24 pp., 16 ill. 

Tests of Significance of Means , J. G. Waugh, Jan. 27, 1945. 9 pp., 2 tables, 1 ill. 


IPC 68 


338 


BIBLIOGRAPHY 


CIT No. 
IPC 69 
Prel. 
IPC 69 
Rev. 
IPC 70 

IPC 71 

IPC 73 

IPC 74 
Prel. 
IPC 74 

IPC 75 

IPC 76 

IPC 77 
Prel. 
IPC 77 

IPC 79 

IPC 80 

IPC 81 

IQC 1 
IQC 2 


IQC 3 
ITC 1 
ITC 2 
ITC 3 
ITC 4 


ITC 5 

IZC 1 

IZC 1 
Rev. 

IZC 2 
IZC 2.1 

IZC 2.2 
IZC 3 
K 3.1 


K 3.3 


LBC 1 
MBE 17 


NDRC No. OSRD No. 

Optical Method, of Measuring Angular Changes of Model Projectiles During Entry 
Phase , B. H. Rule, Jan. 22, 1945. 15 pp., 10 ill. 

2430 Optical Method of Measuring Angular Changes of Projectile Model During Entry 
Phase , B. H. Rule, Mar. 17, 1945. 11 pp., 10 ill. 

Calculated Motion of Forward-Fired Spin-Stabilized Rockets, Including Aerodynamic 
Effects, T. H. Pi, Mar. 2, 1945. 45 pp., 28 ill. 

Computation of Drag and Deceleration Coefficients of 5.0-in. /10 GPSR, T. H. Pi, 
Mar. 12, 1945. 12 pp., 3 tables, 6 ill. 

Computing Manual for Muroc Flight Test Station, Apr. 10, 1945. 34 pp., 5 tables, 

10 ill. 

Shearing Stress on Projectiles at Water Impact, Max Mason and L. B. Slichter, Apr. 
14, 1945. 12 pp., 2 tables, 6 ill. 

2483 Shearing Stress on Projectiles at Water Impact, Max Mason and L. B. Slichter, Apr. 
14, 1945. 11 pp., 2 tables, 6 ill. 

The Production of Instability by the Magnus Moment, L. Davis, Jr., and J. W. Follin, 
Apr. 3, 1945. 15 pp., 5 ill. 

The Effect of Wind on Ground-Fired Spin-Stabilized Rockets during Burning, J. W. 

Follin, Jr., Apr. 23, 1945. 17 pp., 3 tables, 6 ill. 

An Analytical Approximation for the Underwater Trajectory of a Nonrotating Pro- 
jectile, N. A. Haskell, May 28, 1945. 33 pp., 1 table, 14 ill. 

An Analytical Approximation for the Underwater Trajectory of a Nonrotating Pro- 
jectile, N. A. Haskell, June 1, 1945. 36 pp., 14 ill. 

The Determination of a Rocket Trajectory, C. F. Robinson, T. H. Pi, and L. I. Ep- 
stein, July 23, 1945. 39 pp., 9 tables, 18 ill. 

Considerations Involved in the Design of Short-Burning, Long-Range Rockets, L. 
Davis, Jr., Sept. 10, 1945. 19 pp., 4 ill. 

Photographic Studies in Underwater Ballistics, P. M. Hurley, J. S. Fassero, and 
R. C. Jackson, Nov. 15, 1945. 114 pp., 109 ill. 

BR Fragmentation, O. C. Wilson, July 7, 1943. 9 pp., 11 tables, 7 ill. 

Comparison of Fragmentation of the 4.5 -in. Barrage Rocket with the 105-mm Howit- 
zer Shell, O. C. Wilson, C. A. Wirtanen, and J. A. Gilbert, July 30, 1943. 8 pp., 
5 tables, 6 ill. 

Fragmentation Tests on Special BR Bodies, O. C. Wilson, C. A. Wirtanen, and J. A. 
Gilbert, Aug. 19, 1943. 3 pp., 3 tables. 

Yaw and Deflection of UP's Developed During Burning, L. Davis, Jr., Feb. 23, 1942. 
5 pp., 1 table, 1 ill. 

Relative Yaws and Deflections of UP's due to Malalignment for Neutral, Regressive, 
and Progressive Burning, L. Davis, Jr., June 1, 1942. 6 pp., 2 ill. 

The Effect on Dispersion of Variation in the Propellant Temperature for the ASB, 
L. Davis, Jr., June 25, 1942. 4 pp., 1 table. 

Ballistics of Firing an ASB Backwards from an Airplane, L. Davis, Jr., Aug. 10, 
1942. 6 pp., 1 ill. 

Effect of Fins on Yaw and Deflection, L. Blitzer and L. Davis, Jr., Oct. 31, 1942. 
26 pp., 5 tables, 12 ill. 

Style Manual for Reports Published by the Editorial Office, Nov. 29, 1943. 20 pp., 

11 ill. 

Style Guide for Authors of Monographs, Sept. 4, 1945. 11 pp. 

Style Guide for Editors and Printers of Monographs, Nov. 8, 1945. 

Supplement No. 1 to Style Guide for Editors and Printers of Monographs, Dec. 17, 
1945. 

Supplement No. 2 to Style Guide. 

Explosives Safety , B. H. Sage, Nov. 1, 1945. 46 pp., 1 table. 

The CIT Rotating Mirror Camera {Mod 2), I. S. Bowen, Apr. 27, 1945. 21 pp., 2 
tables, 11 ill. 

Auxiliary Equipment for CIT Rotating Mirror Camera and Further Notes on the 
Camera, I. S. Bowen, June 8, 1945. 18 pp., 8 ill. 

Elementary Principles of Rocketry, O. C. Wilson, July 17, 1943. 10 pp., 2 ill. 

The British 3-in. Aircraft Rocket, an Abstract, May 1, 1943. 12 pp., 6 ill. 



NDRC No. OSRD No. 


BIBLIOGRAPHY 


339 


CIT No. 
MBE 18 

MTC 1 

MTC 2 
MTC 3 
MTC 4 
MTC 5 
MTC G 
MTC 7 
NOC 3.1 
NOC 3.2 
NOC 3.3 


NOC 4.1 

NOC 4.2 
NOC 4.3 

NOC 5.1 

NOC 5.2 


NOC 5.3 
NOC 5.4 
NOC 5.5 


NOC 5.6 

NOC 5.7 

NOC 5.8 
and Addenda 
NOC 5.9 

NOC 5.10 

NOC 5.11 


NOC 6.1 
NOC 6.2 

NOC 7.1 

NOC 7.2 


Forward Firing of Rocket Projectiles from British Aircraft, an Abstract, June 16, 1943. 
19 pp. 

The Didion- Bernoulli Approximation for the Trajectory of a Projectile When the Re- 
sistance is Proportional to the Square of the Velocity, L. Davis, Jr., Mar. 23, 1942. 
7 pp., 1 ill. 

Initial Conditions for the Calculation of the CWG Trajectories, L. Davis, Jr., Mar. 
21, 1942. 7 pp., 1 ill. 

Yaw and Deflection of UP’s Developed during Burning, L. Davis, Jr., Mar. 31, 1942. 
7 pp., 4 ill. 

The Theory of the Variation with Temperature of the Dispersion of the CWG, L. Davis, 
Jr., May 1, 1942. 6 pp., 2 tables, 3 ill. 

Effect of Fins on the Yaw and Deflection of CWG’s, L. Blitzer, June 1, 1942. 11 pp., 
2 tables, 6 ill. 

The Effect of the Wind on the Trajectory of a UP, L. Davis, Jr., July 7, 1942. 7 pp., 
1 ill. 

Effect of Fins on the Yaw and Deflection Developed During Burning of UP’s, L. Blitzer 
and L. Davis, Jr., Aug. 10, 1942. 16 pp., 5 tables, 5 ill. 

Naval Ordnance Laboratory Centrifuge Checks of the Calibration of the CIT Midget 
Step Accelerometer, O. D. Terrell, Mar. 15, 1945. 1 p., 1 table. 

Short Duration Acceleration Tests of the CIT Midget Step Accelerometer, O. D. Terrell, 
May 1, 1945. 2 pp., 2 ill. 

The Shock Mounted CIT Step Accelerometer Applied to the Study of Impidsive Vel- 
ocity Changes Associated with the High-Speed Water Entry of Aircraft Torpedoes, 
O. D. Terrell, Aug. 10, 1945. 9 pp., 2 tables. 

Cone Centering Device — Mk XII-1 Gyro, Modifications 1 and 2, R. R. Stokes, 
Feb. 22, 1945. 4 pp., 5 ill. 

Pallet Blade Centering Device, R. R. Stokes, Feb. 26, 1945. 1 p. 

Afterbody and Gyro Pot Pressures of the Mk 13 Torpedo under Hot Steady Running 
Conditions, O. D. Terrell, Mar. 15, 1945. 1 p., 1 table. 

Structural Damage Report on Mk 25-00 No. 5, W. H. Saylor and D. A. Kunz, Mar. 
7, 1945. 6 pp., 3 tables, 2 ill. 

Report on Structural Tests— 176-L Cast Aluminum Afterbody with Cast Aluminum 
Shroud Ring, D. A. Kunz, Mar. 15, 1945. 4 pp., 4 ill. 

Report on Mk 25-00 Assembly and Structural Tests, D. A. Kunz, Mar. 29, 1945. 
5 pp., 3 ill. 

Mk 25 Depth Mechanisms , FBD 1-25 and FJD 1-25, D. A. Kunz, Apr. 2, 1945. 
5 pp., 2 tables, 8 ill. 

Report on Structural Tests — 176-M Cast Aluminum Afterbody ( American Can Co. 
Drwg. EX 1927) with Cast Aluminum Shroud Ring, D. A. Kunz, Apr. 18, 1945. 
4 pp., 4 ill. 

Summary of Information to Date on Mk 13-2, 176 and Westinghouse Joints, D. A. 
Kunz, May 8, 1945. 2 pp., 3 ill. 

Summary of Structural Tests to Date on Cast Aluminum Afterbodies for the Mk 25 
Torpedo, D. A. Kunz, May 8, 1945. 2 pp. 

Structural Tests on Aluminum Afterbodies Cast by CIT — 176-N , D. A. Kunz, May 
29, 1945. 11 pp., 8 ill. (Addenda dated June 9 and 19, 1945.) 

Report on Structural Tests — Mk. 25-00 Propeller Assembly (. American Can Co. Dwg. 

W1-17-L1 ), D. A. Kunz, June 28, 1945. 5 pp., 4 tables, 5 ill. 

Report on Structural Tests — Mk 25-00 Free Wheeling Propeller Assembly Unit No. 1 
(American Can Co. Assembly Dwg. W 1-25-L1 ), D. A. Kunz, Aug. 6, 1945 . 5 pp., 

1 table, 2 ill. 

Report on Structural Tests — Mk 25-00 Free Wheeling Propeller Assembly Unit No. 2 
(American Can Co. Assembly Dwg. W 1-25-L1 ), D. A. Kunz, Aug. 20, 1945. 3 pp., 

2 tables, 2 ill. 

Gyroscopic Orientation Recorder, S. Baker, Sept. 28, 1945. 23 pp., 2 tables, 14 ill. 
Attitude and Roll Study of Mk 13 Shape Torpedo Subsequent to Entry, Saul Baker, 
Jan. 11, 1946. 22 pp., 16 ill. 

Damage to the Special I, CIT Stiffened Afterbody, T. Curtis, Feb. 7, 1945. 6 pp., 
1 table, 7 ill. 

Damage to Mk 26-3 Exercise Head, T. Curtis, Mar. 21, 1945. 3 pp., 5 ill. 



340 


BIBLIOGRAPHY 


CIT No. 
NOC 7.2 
App. 

NOC 11.1 
NOC 15.1 


NOC 17.1 
NOC 20.1 

NOC 24.1 

NOC 32.1 

NOC 32.2 

NOC 32.3 

NOC 35.1 

NOC 35.2 

NOC 38.1 

NOC 42.1 

NOC 43.1 

NOC 44.1 


NOC 45.1 
NOC 46.1 
NOC 47.1 
Rev. 

NOC 47.2 
NOC 48.1 
NOC 48.2 


NOC 50.1 
NOC 51.1 
NOC 53.1 

NOC 53.2 
NOC 53.3 
NOC 53.4 

NOC 53.5 

NOC 53.7 
NOC 53.8 
NOC 55.1 

NOC 56.1 
NOC 57.1 
NOC 58.1 


NDRC No. OSRD No. 

Appendix to NOC 7.2 , 21 Mar., 1945, “ Damage to Mk 26-3 Exercise Head,” D. E. 
Hudson, May 14, 1945. 1 p. 

Propeller Fixity Investigations, T. Curtis, Oct. 2, 1945. 1 p., 1 table. 

The Effect of the Entry Angle Changer on the Underwater Trajectories of Torpedoes 
Launched at the MDH Station, William H. Christie and J. M. French, Sept. 5, 
1945. 5 pp., 1 table, 2 ill. 

Flare Specifications, A. K. Billmeyer, Sept. 30, 1945. 5 pp., 2 ill. 

Preliminary Report on the Mk 2-0 D and R Program ( TLP 20), F. R. Watson, H. 

Bane, and L. Barre, July 10, 1945. 9 pp., 2 tables, 3 ill. 

Time Relationships of Certain Entry Phenomena for the Mk 13 Torpedo at 20° Entry 
Angle, S. Baker and R. Stokes, Sept. 11, 1945. 7 pp., 1 table, 4 ill. 

Design of Torpedo Launching Tube, D. E. Hudson and O. D. Terrell, Mar. 14, 1945. 

9 pp., 7 ill. 

Firing Circuit for the Rocket Booster Stations ( Schematic Diagram TL 3025), R. N. 
Skeeters, July 26, 1945. 3 pp., 1 ill. 

Calculated Performance Characteristics of 36-in. and 49-in. Compressed Air Launch- 
ing Tubes, O. D. Terrell, Aug. 21, 1945. 12 pp., 8 ill. 

Hydrostatic Pressure Test of U Y” and Breech Sections, J. T. Bowen, July 19, 1945. 

10 pp., 8 ill. 

Preliminary Test of “Y” and Breech Sections, J. T. Bowen, June 8, 1945. 6 pp., 
5 ill. 

Variable Angle Torpedo Launcher, Preliminary Study, Francis Carlisle, Aug. 9, 1945. 
21 pp., 11 tables, 1 ill. 

Hydrophone-Signal Amplifier, for Interval Timer, U. E. Younger, Sept. 25, 1945. 

4 pp., 2 ill. 

Drop Table Camera — Operating Instructions, K. Johnson, May 17, 1944, retyped 
Sept. 27, 1945. 2 pp. 

Electrical Installation at Camera Control Station, Drawings TL 3143, 3144 , 3145, 
3146, 3147, 3148, 3149, 3350, 3351, 3379, 3382, R. N. Skeeters, July 27, 1945. 

5 pp. 

Net Data Inconsistencies, H. N. Bane, July 13, 1945. 1 p. 

Cap Firing Mechanism, H. N. Bane, Mar. 26, 1945. 4 pp., 5 ill. 

Mechanism of Pitch Sensitivity for Aircraft Torpedoes, H. Wayland, Feb. 27, 1945. 
10 pp., 7 ill. 

Pitch Sensitivity of Depth of Dive for Mk 13-6 Hot Shots, H. Wayland, May 28, 1945. 
4 pp., 2 ill. 

Velocity-Time and Distance-Time Curves for War Weight Hot Shots, R. W. Ager, 
Apr. 10, 1945. 3 pp.,2 ill. 

Velocity-Distance, Velocity-Time, and Distance-Time Curves for Underwater Travel 
of Various Torpedoes, E. D. Cornelison and J. G. Waugh, Sept. 30, 1945. 10 pp., 
1 table, 8 ill. 

Study of Velocity Measurements, U. E. Younger, Sept. 21, 1945. 5 pp., 3 ill. 

Analysis of Mk 13 Propeller Steel, T. Curtis, Feb. 21, 1945. 1 p. 

Operation of ( 1 ) Midget Neon Lamps , (2) Small B Batteries and {3) Timing Oscilla- 
tors, U. E. Younger, Apr. 27, 1945. 4 pp., 5 tables, 4 ill. 

Regulated Power Supply Circuit, U. E. Younger, May 22, 1945. 6 pp., 4 tables, 4 ill. 
A 100-Cycle Timing Oscillator, U. E. Younger, June 28, 1945. 3 pp., 4 ill. 

Results of Tests on Carbon Strain Gages, U. E. Younger, Dec. 11, 1943, retyped 
July 10, 1945. 2 pp., 2 tables, 1 ill. 

V ibrating-Reed Drive Unit, U. E. Younger, Apr. 3, 1944, retyped July 10, 1945. 
3 pp., 3 ill. 

Electronic Bore-Rider Circuit, U. E. Younger, Sept. 28, 1945. 4 pp., 1 ill. 

Tests on Westinghouse Crystal Accelerometer, A. S. Voak, Sept. 30, 1945. 4 pp., 1 ill. 
High-Density Liquids for Use with Torpedo Exercise Heads, Saul Baker, Oct. 2, 1945. 

6 pp., 1 table, 1 ill. 

Waterproofing of Strain Gages, H. N. Bane, July 9, 1945. 1 p. 

The 35-mm Miller Oscillograph, H. N. Bane, July 13, 1945. 2 pp., 1 table, 5 ill. 

The Determination of Underwater Velocities and Decelerations of Torpedoes by the 
Use of Electrified Nets, William H. Christie and H. N. Bane, Sept. 28, 1945. 6 
pp., 4 ill. 


BIBLIOGRAPHY 


341 


CIT No. 

NOC 59.1 

NOC 60.1 

NOC 61.1 
OBC 3 

OBC 14.2 

OBC 14.3 

OBC 14.4 

OBC 14.5 

OBC 15 

OBC 16 
OBC 18 
OBC 19 
OBC 33 
OBC 41.1 
Oct. 25, 1944 
OBC 41.1 
Aug. 9, 1945 
OBC 41.2 
Aug. 17, 1945 
OBC 43.1 
Nov. 7, 1944 
OBC 49.1 
Jan. 10, 1945 

OBC 49.1 
Sept. 17, 1945 
and att. of 
Sept. 24, 1945 
OBC 52 

OBC 53 
ODC 2 

ODC 5 


ODC 7 

ODC 9 

ODC 10 

OEC 13 
OEC 13.2 

OGC 15 
OGC 24 
OHC 4 
OHC 4.2 
OHC 4.3 

OHC 8 


NDRC No. OSRD No. 

Cavitation and Bubble Studies, William H. Christie and F. C. Watson, Sept. 24, 
1945. 7 pp., 4 ill. 

Steady State Running Altitude and Depth Rudder Position for the Mk 13 Torpedo, 
J. H. Carr and R. W. Haussler, Oct. 1, 1945. 7 pp., 3 ill. 

Neon Tube Cameras, Saul Baker, Oct. 15, 1945. 15 pp., 17 ill. 

Rocket Projectiles Being Developed and Tested at Calif. Inst., Contract OEMsr-250, 
May 5, 1942. 


A-76 

768 

Vertical Antisubmarine Bomb ( VASB ); Vertical Flare (VF), W. A. Fowler, July 23, 
1942. 23 pp., 5 tables, 5 ill. 

A-52M 

872 

Vertical Bombing, C. D. Anderson, W. N. Arnquist, and F. C. Lindvall, Aug. 25 
1942. 10 pp., 2 tables, 5 ill. 

A-54M 

911 

Vertical Bombing, C. D. Anderson, W. N. Arnquist, and F. C. Lindvall, Sept. 12, 
1942. 28 pp., 5 tables, 19 ill. 

A-141 

1242 

Vertical Bombing, Third Report, C. D. Anderson and others, Nov. 25, 1942. 60 pp., 
16 tables, 17 ill. 

A-51M 

805 

Progress on the Installation and Training Program for the Antisubmarine Bomb, or 
“Mousetrap” , W. R. Smythe, Aug. 15, 1942. 7 pp. 


Status of 3-in. AA, F. Roach, Sept. 15, 1942. 1 p. 

BR Crew Training Manual, Nov. 2, 1942. 9 pp. 

Air Resistance of the CWB, L. Davis, Jr., and Hsue-Shen Tsien, Feb. 19, 1943. 3 pp. 
Standardization of Projectiles on Basis of Field Tests, J. Foladare, Apr. 1, 1943. 8 pp. 
3.5-in. and 5-0-in. Spin-Stabilized Rockets, 40 pp., 10 tables, 19 ill. 

Firing Tables for the 5.0-in./5 HCSR Model 134 from the Precise, Mk 50, and Mk 51 
Launchers, P. W. Stoner and J. W. Follin, Jr., 20 pp., 14 tables. 

Range Tables for the 5.0-in./2 HCSR, Model 151-A, with 3.84-lb Grain, P. W. Stoner 
and J. W. Follin, Jr., 7 pp., 5 tables. 

Results of Mallaunching Calculations, L. Davis, Jr., and P. W. Stoner. 18 pp., 1 
table, 8 ill. 

Design and Development of the 11.75-in. Rocket Motor Mark 1 up to October 1, 1944, 
C. W. Snyder, S. Rubin, L. H. Mahony, C. S. Cox, and J. N. McClelland, Jan. 
10, 1945. 88 pp., 75 ill. 

Development of Constant-Burning-Time 11.75-in. Motor for Use as Booster, C. W. 
Snyder, Sept. 17, 1945 and Sept. 24, 1945. 10 pp., 1 table, 4 ill. and 3 pp., 3 
tables. 

Preliminary Sighting Data for 5.0-in. High-Velocity Aircraft Rocket, June 22, 1944. 
8 pp., 7 tables. 

Brief History of the Project, T. Lauritsen, May 11, 1944. 10 pp. 

Microscopic Examination of Extruded Ballistite, W. N. Lacey and B. H. Sage, Nov. 
21, 1941. 9 pp., 6 ill. 

Comparison of Experimental and Predicted Pressure Distribution around a 1.7-in. 
by 0.6-in. by 11.5-in. Grain, R. N. Wimpress, B. H. Sage, and W. N. Lacey, 
Apr. 13, 1942. 10 pp., 2 tables, 4 ill. 

A-40M 404 Remarks on the Use of Double-Base Powder in Jet-Propelled Devices, C. C. Lauritsen, 

Feb. 23, 1942. 12 pp. 

Propellant Grains for Use in a 2.5-in. Motor, Propellants Section, Nov. 5, 1942. 
8 pp., 2 tables, 4 ill. 

Extrusion and Burning Characteristics of a Nitrotoluene Propellant, Propellants 
Section, Nov. 1942. 6 pp., 2 tables, 1 ill. 

Installation of Projectors of PBY, F. C. Lind vail, Sept. 5, 1942. 3 pp., 1 ill. 

Proposed Vertical Bombing Projectors for PBY-5A, F. C. Lindvall, Oct. 14, 1942 
12 pp., 11 ill. 

Partial-Burning Equipment, Nov. 3, 1942. 

A Short Description of the Yaw Machine, Nov. 17, 1943. 1 p. 

Method of Computing External Ballistic Data, Apr. 12, 1943. 4 pp. 

Method of Computing External Ballistic Data, Apr. 19, 1943. 6 pp., 1 table. 

Method of Computing External Ballistic Data, Second Edition, Sept. 30, 1943. 10 pp., 
1 ill. 

Methods of Computing Ballistic Data for Rotating Rockets, Jan. 18, 1944. 7 pp. 


342 


BIBLIOGRAPHY 


CIT No. NDRC No. 
OHC 9 

OIC 3 

OIC 7 
OIC 13 

OMC 18.1 
July 27, 1945; 

Rev. 

Aug. 10, 1945 
OMC 27 
OMC 27.2 
OMC 27.3 
ONC 1 


OPC 3.1 
Sept. 25, 1944 
OPC 3 
Nov. 17, 1944 
OPC 3.1 
Dec. 9, 1944 
OPC 3.1 
Dec. 12, 1944 
OPC 3.1 
Dec. 14, 1944 
OPC 3.1 
Jan. 9, 1945 
OPC 3 
Jan. 10, 1945 
OPC 3 
Jan. 23, 1945 
OPC 3.1 
Apr. 23, 1945 
OPC 3.1 
May 31, 1945 
OPC 3.1 
June 2, 1945 
OPC 3.1 
June 20, 1945 
OPC 6 

OPC 13 

OPC 15 
Dec. 22, 1943 
OPC 15 
Dec. 23, 1943 
OPC 15.1 
May 24, 1944 
OPC 15.1 
Dec. 22, 1944 
OPC 15 
Dec. 26, 1944 
OPC 15.1 
Jan. 10, 1945 
OPC 15 
Jan. 12, 1945 


OSRD No. 

Proposed Airplane Rocket Facilities at Naval Ordnance Test Station, lnyokern, 
G. Mosteller, Mar. 8, 1944. 2 pp., 6 ill. 

Preliminary Report on the HIR Mk III Fuze, V. Rasmussen, Sept. 12, 1942. 4 pp., 
2 ill. 

Type HIR Mark III — Fuze Tests, 6-22 Aug., 1942, B. H. Rule, Sept. 3-19, 1942. 
Assembly and Disassembly of CIT Rocket Fuzes, Special Edition for Rocket School, 
O. C. Wilson 

A Consideration of Vertical Pitch Measurements by Means of Motion Pictures, C. D. 
Anderson and G. M. Safonov. 11 pp., 4 ill. 


Preliminary Data Sheets, Rocket Projectiles, May 16, 1944. 9 pp., 3 ill. 

Preliminary Data Sheets, Rocket Projectiles, June 1, 1944. 32 pp., 19 ill. 

Preliminary Data Sheets, Rocket Projectiles , Aug. 25, 1944. 19 pp., 5 tables, 9 ill. 

A Suggested Method of Scoring to Increase the Marksmanship of Antiaircraft Gun 
Crews and to Decrease the Necessary Training Period, Alex E. S. Green, May 1, 
1942.4 pp. 

Comment on the Damping Moments of a Rocket in Flight , L. Blitzer. 2 pp. 

Tip-Off about a Fixed Point, W. Hayes. 3 pp., 3 ill. 

Effective Burning Times of Rockets for Exterior Ballistic Calculations, L. Blitzer. 

2 pp., 3 ill. 

The Calculation of the Mallaunching Produced by Statically or Dynamically Un- 
balanced Rounds. {Preliminary) , L. Davis, Jr. 9 pp., 1 ill. 

Amplification and Extension of CIT /OPC 2.1, 12 Dec., 1944, L. Davis, Jr. 3 pp., 

1 ill. 

A New System of Coordinates, J. W. Follin, Jr. 2 pp. 

The Inertial Forces on the Parts of a Mallaunched Rocket, L. Davis, Jr. 5 pp., 1 ill. 

Corrections to and an Extension of CIT /OPC 3,10 Jan., 1945 , “ The Inertial Forces 
on the Parts of a Mallaunched Rocket ” L. Davis, Jr. 2 pp. 

Vertical Deviation and Perpendicular Deviation, L. Davis, Jr. 2 pp. 

Trajectories of Rockets at Short Ranges, L. C. Damsgard and L. Davis, Jr. 12 pp., 

3 tables, 3 ill. 

Effective Burning Times and Velocities of Rockets for New Trajectory Tables, R. 

Barrett and L. Blitzer. 14 pp., 15 tables. 

Displacements of the 11.75-in. AR due to Cross Wind Lift and Bernoulli Effect, P. W. 
Stoner and L. Davis, Jr. 

Comparison of Water Tunnel Tests on the 2%-in. Projectile with Field Results, L. 
Davis, Jr., June 12, 1943. 3 pp. 

Effect of Launcher Length on the Downward Deflection of the Trajectory of the 4-5-in. 

BR {4H-01 Ma, Lot If) at End of Burning, J. G. Waugh. 3 pp., 1 ill. 

Qualitative Discussion of Equilibrium Yaw, S. Rubin. 4 pp., 2 ill. 

Experimental Measurement of Yaw Moment Coefficient, S. Rubin. 2 pp. 

Test of the Significance of Differences in Mean Deviations from the Centroid of Rocket 
Groups, P. H. Taylor. 6 pp., 2 tables, 3 ill. 

A Theoretical Analysis of the Results of the Field Firing of Dynamically Unbalanced 
3.5-in SSR Model 123 on 22 Nov., 1944, J- W. Follin, Jr. 8 pp., 3 tables, 1 ill. 

An Approximate Formula for the Variation of Dispersion due to Dynamic Unbalance 
with Launcher Length and Burning Time, L. Davis, Jr. 4 pp., 1 ill. 

External Ballistics of 5 -in. HVSR Fired Rearward from Aircraft, G. Safonov. 4 pp., 

2 tables, 1 ill. 

Approximate Curves for the Trajectory of the 5.0-in HCSR during the First 8 sec, 
L. Davis, Jr. and L. I. Epstein. 3 pp., 1 ill. 


BIBLIOGRAPHY 


343 


CIT No. 

OPC 15 
Aug. 2, 1945 
OPC 15 
Feb. 12, 1945 
OPC 15.1 
Jan. 18, 1945 
OPC 15.1 
Feb. 15, 1945 
OPC 15.1 
Mar. 6, 1945 
OPC 15.1 
Apr. 5, 1945 
OPC 15.1 
Apr. 30, 1945 

OPC 15.1 
May 12, 1945 
OPC 15.1 
May 26, 1945 

OPC 15.1 
June 19, 1945 
OPC 15 
Oct. 10, 1945 
OPC 15 
Oct. 31, 1945 
OPC 15.2 
Sept. 10, 1945 
OPC 18 
OPC 20 
OPC 21 


OPC 21.2 
OPC 21.3 


OPC 22 

OPC 23 

OPC 24 

OPC 24.2 

OPC 28 

OPC 29 

OPC 29 
Sup.l 

OPC 29 
Sup. 2 
OPC 30 


OPC 32 


NDRC No. OSRD No. 

Supplement to CIT /OPC 15, 12 Jan., 1945, L. Davis, Jr., and L. I. Epstein. 2 pp., 

1 ill. 

The Stability Factors and Equilibrium Yaws of a Number of SSR. Preliminary Report, 
L. Davis, Jr., and L. I. Epstein. 12 pp., 4 tables, 1 ill. 

Closed Breech Ballistic Data, A. B. Meinel. 7 pp., 6 ill. 

Ballistic Data for the 5.0-in./ 5 SmSR Mdl 127, 14 Feb., 1945, P. II. Taylor. 6 pp., 

2 tables, 3 ill. 

Effect of Wind during Burning, J. W. Follin. 6 pp., 1 table. 

Calculation of the Mallaunching and Dispersion Produced by an Uncentered Grain 
in the 5.0-in/ 14 GASR, L. S. Abrams. 6 pp., 1 ill. 

Tentative Values for the Stability Factor and Equilibrium Yaw of the 5.0-in. /I HCSR 
Model 50. Preliminary Report. Supplement No. 1 to CIT /OPC 15, 12 Feb., 1945, 
L. I. Epstein. 2 pp. 

Minimum Clearance Angle for 5.0-in./5 HCSR, L. Davis, Jr., and J. W. Follin, Jr. 

3 pp., 1 table, 1 ill. 

Tentative Values for the Stability Factor and Equilibrium Yaw of the 5.0-in./ 2 HCSR 
Model 51; Preliminary Report. Supplement No. 2 to CIT /OPC 15, 12 Feb., 19/5, 
L. I. Epstein. 1 p. 

Terminal Ballistics of Spin-Stabilized Rockets, L. Davis, Jr., 2 pp. 

Mallaunching due to Elliptical Bourrelets, W. B. Dayton. 7 pp. 

Another Application of Yaw Camera Records, W. B. Dayton. 4 pp. 

Determination of Ratio of Effective Burning Time to Reaction Time from Static Burn- 
ing Curves, L. Blitzer. 7 pp., 1 table, 3 ill. 

Effect of Fins on Dispersion of 5.0-in. HVAR, L. Blitzer, May 16, 1944. 5 pp., 3 ill. 
Sight Setting for 5.0-in. HVAR ( P47D ). 7 pp. 

The Effect of Acceleration, Automatic Sights, and Head Wind on the Curve of Ap- 
proach and on the Sight Setting for the Forward Firing of Rockets, L. Davis, Jr. 
11 pp., 1 ill. 

A Correction and an Illustrative Curve for CIT /OPC 21, L. Davis, Jr., and P. W. 
Stoner, Oct. 30, 1944. 3 pp. 

The Effect of Acceleration, Automatic Sights, and Head Wind on the Curve of Ap- 
proach and on the Sight Setting for the Forward Firing of Rockets (Revision of 
CIT/OPC 21 and CIT/OPC 21.2), L. Davis, Jr., Nov. 24, 1944. 13 pp., 2 ill. 
Effective Angle of Attack Data, TBF-1C, TBM-1C, F4U-1 , F4U-1D, FG-1, FM-2, 
F6F-3, F6F-5, Oct. 6, 1944. 5 pp. of tables. 

Gravity Drops (mils) Normal to Effective Launching Line for 2.25-in. AR Model 1; 

Mk 5 Launcher, revised July 15, 1944. 6 pp., all tables. 

The Determination of the Angle of Attack in a Vertical Plane, L. Davis, Jr., Oct. 30, 
1944. 8 pp., 1 table. 

The Determination of the Angle of Attack in a Vertical Plane, L. Davis, Jr., Nov. 24, 
1944. 9 pp., 1 table. 

Effect of Small Changes in Burning Time and Velocity on Trajectory Drops of Rockets 
Fired from Aircraft, L. Blitzer, Nov. 1, 1944. 7 pp., 4 tables. 

Motion of the 5.0-in. Hybrid SSR in Forward Firing from Aircraft as Inferred from 
Yaw Camera Records, L. I. Epstein, Feb. 17, 1945. 11 pp., 6 ill. 

Second Report on the Motion of the 5 .0-in./10 Hybrid GASR in Forward Firing from 
Aircraft as Inferred from Yaw Camera Records, L. I. Epstein, Feb. 17, 1945. 
5 pp., 2 ill. 

Motion of the 5 .0-in/14 GASR Model 39 in Forward Firing as Inferred from Yaw 
Camera Records, L. I. Epstein, May 26, 1945. 2 pp. 

The Determination of the Stability Factor and the Ratio of the Moments of Inertia of a 
Spinner from Yaw Camera Measurements, L. Davis, Jr., Feb. 26, 1945. 6 pp., 
1 ill. 

The Stability Factor Deduced from the Yaw Camera Record and the Effect on it of the 
Magnus Moment, P. W. Stoner and L. Davis, Jr., Apr. 4, 1945. 5 pp., 2 tables. 



BIBLIOGRAPHY 


344 


CIT No. NDRC No. 

OSRD No. 


OPC 33 


A Note on the Reasons Why the Same Spin-Stabilized Rocket Cannot be Used for Very 
Accurate Fire with a Flat Trajectory and for Barrage Purposes , L. Davis, Jr., 
Apr. 4, 1945. 4 pp., 1 table. 

OPC 34 


Curves for Calculation of Trajectories during Burning of Fin-Stabilized Rockets, L. 
Blitzer, L. I. Epstein, and M. C. Houghton, May 3, 1945. 15 pp., 8 ill. 

OZC 23 


A System of Notation for Use in Treatments of Exterior Ballistics of Rockets, Theo- 

July 13, 1945 


retical Research Group. 12 pp. 

RBC 1 


Artillery Rockets, C. C. Lauritsen, Mar. 20, 1942. 13 pp., 4 ill. 

RDC 1 


Reprint: Studies of the Jet Propulsion of Submerged Projectiles, B. H. Sage, June 1, 
1942. 155 pp.,6 tables, 17 ill. 

TAC 1 A-149 

1260 

The Dependence of the Masses of Rocket Components on Their Dimensions, L. Davis, 
Jr., Dec. 14, 1942. 21 pp., 3 tables, 4 ill. 

TPC 1 


Calculations of Probable Error, W. R. Smythe, Dec. 8, 1942. 14 pp., 1 table, 3 ill. 

TPC 2 


The Relationship between Dispersion in Firing from a Plane and from the Ground, 
I. S. Bowen, Dec. 25, 1942. 5 pp. 

UAC 1 


Motor Catalogue, Sept. 28, 1942. 12 pp., 7 tables, 3 ill. 

UAC 1.2 


Motor Catalogue Supplementary Internal Ballistics Curves, Nov. 20, 1942. 22 pp., 
22 ill. 

UAC 1.3 


Motor Catalogue: Standardization Curves: Effective Gas Velocity, Maximum Pressure, 
Maximum Pressure Drop , and Burning Time as Functions of Temperature, Feb. 10, 
1943. 11 pp., llill, 

UAC 2 


Internal Ballistic Curves for Neutral-Burning Grains, C. T. Elvey, June 8, 1943. 
23 pp., 22 ill. 

UBC 1 


CIT Rockets, Nov. 18, 1942. 27 pp., 18 tables. 

UBC 3 


Rocket Projectiles, California Institute of Technology, Mar. 1, 1943. 1 p., 1 table. 

UBC 4 


Dispersion Data, Series of Tables for All Projectiles, Apr. 5, 1943. 8 pp., 8 tables. 

UBC 18 


Short Description of Service Rockets, Apr. 15, 1944. 33 pp., 31 ill. 

UBC 27 

2225 

Trajectories of Aircraft Rockets, 3.5-in. and 5.0-in., Sept. 25, 1944. 99 pp. of tables. 

UBC 28 

2409 

Internal and External Ballistic Data, Fin-Stabilized Rockets, Mar. 15, 1945. 325 pp., 
numerous ill. 

UBC 29 

2242 

Trajectories of Aircraft Rockets, 2.25-in. Target, Oct. 10, 1944. 11 pp. of tables. 

UBC 29.2 

2312 

Trajectories of Aircraft Rockets, 2.25-in. Target, Dec. 14, 1944. 11 pp., all tables. 

UBC 30 

2290 

Trajectories of 11.75-in. Aircraft Rockets, Nov. 17, 1944. 17 pp., all tables. 

UBC 32 

2314 

Trajectories of 5.0-in. High-Velocity Aircraft Rocket and 2.25-in. Practice Round. 

C Airplane Speeds in Miles per Hour), Dec. 15, 1944. 26 pp., all tables. 

UBC 34 

2415 

Abridged Catalog, CIT Rockets, Mar. 15, 1945. 122 pp., 6 tables, numerous ill. 

UBC 35 

2540 

Trajectories of Aircraft Rockets, Jan. 4, 1946. 76 pp., 69 tables, 7 ill. 

UEC 2 


Illustrations of Landing Operations Assisted by Barrage Rockets, R. W. Porter, Mar. 
20, 1943. 21 pp., 21 ill. 

UEC 5 

2440 

Abridged Catalog, Rocket Launchers , May 1, 1945. 79 pp., 55 ill. 

UFC 1 


Results of Photographic Measurements of 4.5-in. Rockets on 13-14 Feb., 1943, I. S. 
Bowen, Mar. 2, 1943. 94 pp., 2 tables, 87 ill. 

UIC 1 


Fuze Catalogue, Oct. 1942. 22 pp., 9 ill. 

UIC 2 


CIT Fuze Catalogue, Fuze Group, Jan. 15, 1943. 18 pp., 8 ill. 

UIC 3 


Catalogue; CIT Rocket Fuzes, CIT Fuze Group, Jan. 15, 1944. 68 pp., 2 tables, 
44 ill. 

UIC 3 


Fuzes for U. S. Navy 5.0-in. HVAR {Excerpts from CIT Fuze Catalogue), June 23, 

(Excerpts) 


1944. 16 pp., 9 ill. 

UMC 1 


Bibliography of Published Reports 19 Oct.. 1941 to 27 Jan., 1943. 12 pp. Addenda, 
UMC 1.1 to 1.7, through July 31, 1943, 1 or 2 pp., each. (All obsolete.) 

UMC 4 


Index to Progress Reports, Mar. 7, 1943. (Obsolete.) 

UMC 4.2 


Index to CIT Progress Reports, Oct. 26, 1941 through Aug. 22, 1943. (Obsolete.) 

UMC 4.3 


Index to CIT Progress Reports, Oct. 26, 1941 through May 21, 1944. 

Sup. 1 


Through July 23, 1944. 

Sup. 2 


Through Nov. 12, 1944. 

Sup. 3 


Through Apr. 1, 1945. 

Sup. 4 

2543 

Through Sept. 15, 1945; completes index. 

UMC 5x 


Bibliography of Reports Relating to A$ Warfare, July 2, 1943. 3 pp. 

UMC 7 


Abstract of British Reports on Forward Firing from Aircraft, Aug. 4, 1943. 46 pp., 
21 tables, 27 ill. 


BIBLIOGRAPHY 


345 


CIT No. 

UMC 7.2 

NDRC No. OSRD N< 

UMC 8 
UMC 8.2 
UMC 8.3 

2515 

UMC 8.4 


UMC 42 

UMC 42.2 

2224 

UNC 1 

2118 

UNC 1 

Sup. 1 

UNC 2 

2254 

UNC 3 

2264 

UNC 4 

2271 

UNC 5 

2272 

UNC 6 

2273 

UNC 7 

2274 

UNC 8 

2275 

UPC 1 


UPC 2 


UPC 3 


UPC 4 

2112 

UZC 2 



UZC 3 


UZC 4 


UZC 5 


UZC 6 


Abstract of British Reports on Forward Firing from Aircraft. Supplement to UMC 7, 
Sept. 28, 1943. 28 pp., 17 tables, 6 ill. 

Bibliography of Published Reports , Oct. 19, 1941 to July 31, 1943. 20 pp. (Obsolete.) 

Bibliography of Published Reports, Oct. 19, 1941 to Mar. 16, 1944. 29 pp. (Obsolete.) 

Bibliography of Published Reports, Oct. 19, 1941 to June 30, 1945. (Obsolete with 
issue of this final bibliography, UMC 8.4.) 

Bibliography of Published Reports, Oct. 19, 1941 to Aug. 31, 1946. (Does not include 
9 Final Report Volumes, 2 Monographs, Fowler Manuscript.) 

Abridged Catalogue: Service Rockets, Fuzes, and Launchers, Mar. 15, 1944. 203 pp. 

Supplement to Abridged Catalogue; Rockets, Fuzes, and Launchers, Oct. 10, 1944. 

53 pp., 37 ill. 

Manual for the Computation of Effective Rocket Temperatures, Aircraft Rockets with 
3.25-in. Motors, L. Davis, Jr., June 10, 1944. 68 pp., 30 tables. 

Computation of Effective Rocket Temperatures in 3.25 in., 5.0 in . , and 1 1 .75 in. Motors, 
L. Davis, Jr., July 2, 1945. 

Boresighting and Effective Angle of Attack Data for Various Aircraft, Oct. 25, 1944. 

54 pp., 8 tables, 30 ill. 

FM-2, Sight Settings for 2.25-in., 3.5-in., and 5.0-in. Aircraft Rockets , Nov. 11, 1944. 
77 pp., mostly tables, 20 ill. 

FJffJ-1, FJfU-lD , FG-1, Sight Settings for 2.25-in., 3.5-in., and 5.0-in. Aircraft 
Rockets, Nov. 14, 1944. 77 pp., mostly tables, 18 ill. 

F6F-3, F6F-5, Sight Settings for 2.25-in., 3.5-in., and 5.0-in. Aircraft Rockets, 
Nov. 18, 1944. 77 pp., mostly tables, 19 ill. 

TBM-1, TBF-1, Sight Settings for 2.25-in., 3.5-in., and 5.0-in. Aircraft Rockets, 
Nov. 28, 1944. 77 pp., mostly tables, 12 ill. 

TBM-1C, TBF-1C, Sight Settings for 2.25-in., 3.5-in., and 5.0-in. Aircraft Rockets, 
Dec. 1, 1944. 77 pp., mostly tables, 13 ill. 

SB2C-1 , SB2C-1C, SB2C-3, SB2C-4, Sight Settings for 2.25-in., 3.5-in., and 5.0-in. 
Aircraft Rockets, Nov. 23, 1944. 77 pp., mostly tables, 19 ill. 

Ranges and Velocities as Functions of Propellant and Projectile Weights, June 1943. 
6 pp., 1 table, 5 ill. 

Determination of Sight Settings: 3.5-in. and 5.0-in. Aircraft Rockets, Mk 4 and 
Mk 5 Aircraft Launchers, Mar. 20, 1944. 59 pp. 

Determination of Sight Settings, 3.5-in. and 5.0-in. Aircraft Rockets from Zero- 
Length Launcher, Army Aircraft, Apr. 25, 1944. 29 pp. 

Gravity Drop Tables, British 3.25-in. Rocket Motor No. 1 Mk 2 with American 3.25- 
in. and 5.0-in. Bodies, May 25, 1944. 21 pp., 20 tables. 

Report on the Transfer of the Goldstone Field Testing Station, Section C of Contract 
OEMsr-418 from California Institute of Technology Operating Under the Office of 
Scientific Research and Development to the U. S. Navy at U. S. Naval Ordnance 
Test Station, Inyokern, California, Oct. 15, 1945. 82 pp. 

Report on the Transfer of the Morris Dam Hydrodynamics Station, Section VII of 
Contract OEMsr-418 from California Institute of Technology Operating Under the 
Office of Scientific Research and Development to the U . S. Navy, U. S. Naval Ord- 
nance Test Station, Inyokern, California, Oct. 15, 1945. 86 pp. 

Report on the Transfer of the Rocket Developmental Engineering Section of Contract 
OEMsr-418 from California Institute of Technology Operating Under the Office of 
Scientific Research and Development to the U . S. Navy ( Navy Contracts NOrd-9286 
and NOrd ( F ) — 1428), July 31, 1945. 374 pp. 

Volume I, Report on the Transfer of the Foothill Camel Plant Section of Contract 
OEMsr-418 from California Institute of Technology Operating Under the Office of 
Scientific Research and Development to the U. S. Navy {Navy Contracts NOrd-9286 
and NOrd ( F)—1428 ), Oct. 15, 1945. 15 pp. 

Volume II, Capital Equipment Inventory Transferred from OSRD to the Navy Con- 
tract OEMsr-418 Foothill Camel Plant Section California Institute of Technology, 
Pasadena, California. Transferred to'.Bureauof Ordnance Contract NOrd (F) — 1428, 
Oct. 15, 1945. 188 pp. 


346 


BIBLIOGRAPHY 


CIT No. 
UZC 7 


UZC 8 


NDRC No. OSRD No. 

Transfer from Office of Scientific Research and Development, California Institute of 
Technology Contract OEMsr-418, Section V, to: U. S. Navy, NOTS, Inyokern, 
California, Volume III . Capital Equipment Inventory Transferred from: Eaton 
Canyon Facility, Pasadena, California, Oct. 31, 1945. 134 pp. 

Transfer from Office of Scientific Research and Development, California Institute of 
Technology Contract OEMsr-^18, Section V, to: U. S. Navy, NOTS, Inyokern, 
California, Volume II. Capital Equipment Inventory China Lake and Salt Wells 
Pilot Plants, Oct. 31, 1945. 593 pp. 


v 


BIBLIOGRAPHY FOR VOLUME 


Numbers such as Div. 3-713-M5 indicate that the document listed has been microfilmed and that its title appears 
in the microfilm index printed in a separate volume. For access to the index volume and to the microfilm, consult the 
Army or Navy agency listed on the reverse of the half-title page. 


PART I 
Chapter 1 

1. Water Entry and Underwater Ballistics of Pro- 

jectiles (Final Report), Edited by Alice G. Ritter, 
OSRD 2551, OEMsr-418, California Institute of 
Technology, 1946. Div. 3-713-M5 

2. Photographic Studies in Underwater Ballistics, 

P. M. Hurley, J. S. Fassero, and R. C. Jackson, 
OEMsr-418, Report IPC 81, California Institute of 
Technology, Nov. 15, 1945. Div. 3-722. 2-M2 

Chapter 2 

1. Aircraft Torpedo Development and Water Entry 
Ballistics (Final Report), OSRD 2550, OEMsr-418, 
California Institute of Technology, 1946. 


PART II 
Introduction 

1. The Interior Ballistics of Rockets, R. N. Wimpress, 
monograph (unclassified) to be published by the 
McGraw-Hill Book Company, Inc. 

2. Processing of Rocket Propellants (Final Report), 

W. H. Corcoran and Quentin Elliott, OSRD 2552, 
OEMsr-418, California Institute of Technology, 
1946. Div. 3-300-M2 

Chapter 5 

1. Investigations on the Burning Characteristics of 
Propellant Powder and Their Effects Upon Steady- 
State Pressure in Rocket Motors, Bruce H. Sage, 
OEMsr-418, Report Ms-821/JDC 84, California 
Institute of Technology, Nov. 1, 1945. 

Div. 3-355-M4 

2. Ballistic and Physical Characteristics of a Japanese 

Rocket Propellant, Bruce H. Sage, OEMsr-418, Re- 
port Ms-837/JDC 91, California Institute of Tech- 
nology, Dec. 30, 1945. Div. 3-310-M5 

3. Effect of Opacity on the Burning Characteristics of 

Extruded Ballistite Grains, Bruce H. Sage, OSRD 
2134, OEMsr-418, Report JDC 61, Service Project 
NO-33, California Institute of Technology, Apr. 21, 
1944. Div. 3-361. 514-M10 

4. Some Physical Properties of Double-Base Powders, 
Bruce H. Sage, OEMsr-418, Report JDC 51, Cali- 
fornia Institute of Technology, Oct. 12, 1943. 

Div. 3-361.2-M2 


5. Some Studies of the Physical Properties of Bal- 

listite, Donald S. Clark, OEMsr-418, Report JDC 
36, California Institute of Technology, Feb. 11, 
1943. Div. 3-361. 51-M5 

6. Stabilization of Reaction of Tubular Propellant 

Grains by the Use of Longitudinal Ridges in the 
Central Perforations, Bruce H. Sage, OSRD 2541, 
OEMsr-418, Report JDC 75, Service Projects OD- 
14 and NO-33, California Institute of Technology, 
May 19, 1945. Div. 3-361.524-M6 

7. The Compressive Characteristics of Several Pro- 

pellants Determined at a Constant Rate of Stress 
Application, Donald S. Clark, OSRD 2534, OEMsr- 
418, Report JDC 80, California Institute of Tech- 
nology, Oct. 11, 1945. Div. 3-353-MI 

8. The Use of the Ultimate Strength as Determined 
in a Simple Compression Test as a Measure of JP 
Propellant Quality for Mk 13 Grains, Bruce H. 
Sage, OEMsr-418, Report Ms-834/JDC 89, Cali- 
fornia Institute of Technology, Dec. 1, 1945. 

Div. 3-353-M2 

9. Impact Characteristics of Several Double-Base Pro- 

pellants, Bruce H. Sage, OEMsr-418, Report Ms- 
836/JDC 86, California Institute of Technology, 
Dec. 10, 1945. Div. 3-361.25-MI 

10. Free and Restricted Column Behavior of Some 

Double-Base Propellants, Bruce H. Sage, OEMsr- 
418, Report Ms-838/JDC 92, California Institute of 
Technology, Dec. 30, 1945. Div. 3-361. 26-M6 

11. The Relation of Column Strength to the Ballistic 

Performance of Mk 13 Grains, Bruce H. Sage, 
OEMsr-418, Report JDC 87, California Institute of 
Technology, Dec. 15, 1945. Div. 3-322-M4 

12. Design of a Cruciform Charge for the 3.25-in. 

Motor, Bruce H. Sage, OEMsr-418, Report JDC 
46, California Institute of Technology, July 19, 
1943. Div. 3-322-MI 

13. Investigation of JPH Propellant Lots FDAP 28 

and FDAP 29, Bruce H. Sage, OSRD 2517, OEMsr- 
418, Report JDC 74, Service Projects OD-14 and 
NO-33, California Institute of Technology, June 5, 
1945. Div. 3-330-M3 

14. Development of a 2A-lb Cruciform Charge for the 

5.0-in. Rocket Motor, Bruce H. Sage, OSRD 2108, 
OEMsr-418, Report JDC 62, California Institute of 
Technology, May 4, 1944. Div. 3-322-M3 

15. Static Firing Tests on Large-Diameter Grains of 

Extruded Ballistite, Bruce H. Sage, OEMsr-418, 
Report JDC 12, California Institute of Technology, 
July 30, 1942. Div. 3-361.514-M2 


348 


BIBLIOGRAPHY 


16. Testing of Quality of Small Grains of Extruded 
Ballistite, Bruce H. Sage, OSRD 896, NDRC A-94, 
OEMsr-418, Report JDC 17, California Institute 
of Technology, Aug. 20, 1942. Div. 3-361. 514-M4 

17. Burning Rate of Four-Spoke Grains of Extruded 

Ballistite, Bruce H. Sage, OEMsr-418, Report JDC 
18, California Institute of Technology, Sept. 25, 
1942. Div. 3-361. 514-M6 

18. Influence of Sizes of the Axial Perforation upon 

the Performance of Radially -Burning Propellant 
Grains, Bruce H. Sage, OSRD 966, NDRC A-106, 
OEMsr-418, Report JDC 19, California Institute of 
Technology, Oct. 28, 1942. Div. 3-361.514-M7 

19. Internal Ballistics of Jet-Propelled Devices, Bruce 

H. Sage, OSRD 1069, NDRC A-115, OEMsr-418, 
Report JAC 2, California Institute of Technology, 
Oct. 23, 1942. Div. 3-840-M3 

20. Evaluation of Pressure-Time Relationships Occur- 
ring in Static Firing of Rocket Motors, Bruce H. 
Sage, OSRD 2484, OEMsr-418, Report JGC 8, 
California Institute of Technology, Apr. 10, 1945. 

Div. 3-626-M4 

21. Some Calculations and Experimental Measure- 
ments upon the Pressure Distribution around Thin- 
Webbed Charges During Firing, R. N. Wimpress, 

G. W. Miller, Bruce H. Sage, and William N. 
Lacey, OEMsr-250, Report IDC 10, California 
Institute of Technology, Apr. 8, 1942. 

Div. 3-355-MI 

22. Thermodynamic Properties of Products of Re- 
action of Ballistite (Intermediate Report), Bruce 

H. Sage and William N. Lacey, OSRD 495, OEMsr- 

418, Report JDC 4, California Institute of Tech- 
nology, Feb. 4, 1942. Div. 3-361.51-M2 

23. Development of the Mk 13 Cruciform Propellant 
Grain, Bruce H. Sage, OEMsr-418, Report JDC 56, 
Service Projects OD-26 and NO-33, California In- 
stitute of Technology, Dec. 29, 1943. 

Div. 3-322-M2 

24. Rate of Diffusion of Nitroglycerin through Cellu- 

lose Acetate, Bruce H. Sage, OSRD 2365, OEMsr- 
418, Report JDC 68, Service Projects OD-14 and 
NO-33, California Institute of Technology, Jan. 1, 
1945. Div. 3-361. 214-M5 

25. Development of a Hexoform Ballistite Propellant 

Grain for an 8-in. Rocket Motor, Bruce H. Sage, 
OEMsr-418, Report JDC 93, California Institute of 
Technology, Nov. 15, 1945. Div. 3-323-MI 

26. Development of a J et-Propulsion Unit for the Mk 13 
Torpedo, Bruce H. Sage, OSRD 2539, OEMsr-418, 
Report JOC 4, Service Project NO-177, California 
Institute of Technology, Sept. 12, 1945. 

Div. 3-722.3-MI 


27. Comparison of Experimental and Predicted Pres- 
sure Distribution around a 1.7-in. by 0.6-in. by 
11.5-in. Grain, R. N. Wimpress, Bruce H. Sage, 
and William N. Lacey, OEMsr-418, Report ODC 5, 
California Institute of Technology, Apr. 13, 1942. 

Chapter 6 

1. The Use of Ballistite Turnings in Primers (Pre- 

liminary Report), Bruce H. Sage and William N. 
Lacey, OSRD 577, NDRC A-56, OEMsr-250, Report 
JCC 2, California Institute of Technology, May 14, 
1942. Div. 3-361.53-MI 

2. Preliminary Investigation of Metal-Oxidant Ignit- 

ers for Ballistite, Bruce H. Sage, OEMsr-418, Re- 
port JCC 6, California Institute of Technology, 
Feb. 25, 1943. Div. 3-420-M3 

3. Effect of Squib Boosters on the Performance of 

Black Powder Igniters, Bruce H. Sage, OEMsr-418, 
Report JCC 9, California Institute of Technology, 
Aug. 14, 1943. Div. 3-421-MI 

4. Threaded-Closure Plastic-Case Igniters for 2.25-in. 

Rocket Motors, Bruce H. Sage, OEMsr-418, Report 
JCC 11, California Institute of Technology, Mar. 
16, 1944. Div. 3-422.1-M6 

5. Effect of Relative Humidity on the Water Content 

of Black Powder, Bruce H. Sage, OEMsr-418, Re- 
port JCC 7, California Institute of Technology, 
May 5, 1943. Div. 3-362-MI 

6. Induction Firing for Rockets (Final Report), C. F. 

Bjork and M. Bondy, OSRD 5814, OEMsr-273, ABL 
Report W-18.3, George Washington University, De- 
cember 1945. f . Div. 3-222-M2 

7. Investigation of the Use of Plastic-Case Igniters 
for the ASPC Motor, Bruce H. Sage, OSRD 1318, 
NDRC A-158, OEMsr-418, Report JCC 5, Cali- 
fornia Institute of Technology, Jan. 7, 1943. 

Div. 3-422.1-M2 

8. Development of Cellulose Acetate Igniter Cases 
for 1.25-in. and 2.25-in. Rocket Motors, Bruce H. 
Sage, OEMsr-418, Report JCC 8, California Insti- 
tute of Technology, Aug. 12, 1943. Div. 3-422. 1-M5 

9. Development of Tin-Plate Case Igniters for Artil- 

lery Rockets, Bruce H. Sage, OEMsr-418, Report 
JCC 12, California Institute of Technology, Dec. 30, 
1945. Div. 3-422.2-MI 

Chapter 7 

1. The Extrusion of UP Propellant, 15 /16-in. Solvent- 

less Extruded Ballistite, Thomas L. Lauritsen, 
OEMsr-418, Report JDC 1, California Institute of 
Technology, Dec. 15, 1941. Div. 3-342-MI 

2. Memorandum on Extrusion of Ballistite Tube, Re- 
port MDC 1. 


BIBLIOGRAPHY 


349 


3. Extrusion of Large Tubular Grains of Ballistite, 

Bruce H. Sage, OSRD 1183, NDRC A-135, OEMsr- 
418, Report JDC 24, California Insitute of Tech- 
nology, Dec. 1, 1942. Div. 3-361. 524-MI 

4. Description and Tentative Specifications of a Pro- 
pellants Extrusion Plant for the United States 
Navy. (California Institute of Technology Chemi- 
cal Engineering MS 735. A report under Navy 
contract NORD 500.) 

5. Extrusion of Multi-Web Grains of Ballistite , Bruce 

H. Sage, OSRD 1403, NDRC A-176, OEMsr-418, 
Report JDC 38, California Institute of Technology, 
Feb. 18, 1943. Div. 3-361.524-M4 

6. Experimental Production Facilities at the Eaton 
Canyon Site, Bruce H. Sage, OEMsr-418, Report 
IDC 22, California Institute of Technology, Nov. 
24, 1942. 

7. Propellant Processing, Igniter Construction, and 
Motor Loading Facilities as of January 1, 19 A3, 
Bruce H. Sage, OEMsr-418, Report JDC 45, Cali- 
fornia Institute of Technology, Nov. 24, 1943. 

Div. 3-361.1-M7 

8. A 12-in. Vertical Press for Extrusion of Ballistite, 
Bruce H. Sage, OEMsr-418, Report JDC 53, Cali- 
fornia Institute of Technology, Oct. 26, 1943. 

Div. 3-361. 521-M3 

9. Description of Facilities at the China Lake Pilot 

Plant, Bruce H. Sage, OSRD 2553, OEMsr-418, Re- 
port JDC 83, California Institute of Technology, 
Mar. 13, 1946. Div. 3-610-M6 

10. Ignition of the Powder Charge in an Extrusion 
Press, Bruce H. Sage, OEMsr-418, Report IDC 33, 
California Institute of Technology, May 5, 1943. 


PART III 
Introduction 

1. Rocket Fundamentals, OSRD 3992, OEMsr-273, 
ABL Special Report 4, George Washington Univer- 
sity, 1944. Div. 3-210-M3 

Chapter 8 

1. The Development of the T-A Powder Charge for the 
2.36-in. Rocket Grenade (Final Report), Rufus 
Lumry and L. N. Streff, OSRD 5589, ABL Report 
W-3.4, George Washington University, Nov. 29, 
1945. Div. 3-320-M4 

Chapter 9 

1. Rocket Fundamentals, OSRD 3992, OEMsr-273, 
ABL Special Report 4, George Washington Univer- 
sity, 1944. Div. 3-210-M3 


2. The Reduced Specific Impulse of Ideal Gases (Final 
Report), Nancy Marmer and F. T. McClure, OSRD 
5828, OEMsr-273, ABL Report P-3, George Wash- 
ington University, January 1946. Div. 3-244-M3 

3. Thermodynamic Properties of Propellant Gases, 

Joseph O. Hirschf elder, F. T. McClure, C. F. Cur- 
tiss, and D. W. Osborne, OSRD 1087, NDRC A-116, 
November 25, 1942. Div. 3-350-MI 

4. Thermodynamic Properties of British Flashless and 

Cordite MD Powders, F. T. McClure, D. W. Os- 
borne, and Joseph O. Hirschfelder, OSRD 817, 
NDRC A-82, Carnegie Institution of Washington, 
August 24, 1942. Div. 3-310-MI 

5. Thermodynamic Properties of Special Double-Base 
Powders, D. W. Osborne, F. T. McClure, and Joseph 
O. Hirschfelder, OSRD 1014, NDRC A-107, George 
Washington University, Nov. 16, 1942. 

Div. 3-361.2-MI 

6. Simple Calculation of Thermo chemical Properties 
for Use in Ballistics, Joseph O. Hirschfelder and 
Jack Sherman, OSRD 935, NDRC A-101, October 
1942. 

7. Simple Calculation of Thermochemical Properties 
for Use in Ballistics. Addenda to NDRC Report 
A-101, OSRD 935 (Memoranda A-67M to A-70M), 
Joseph O. Hirschfelder and Jack Sherman, OEMsr- 
51, Service Projects OD-52 and NO-23, Carnegie 
Institution of Washington, March 1943. 

Div. 3-361.51-M6 


Chapter 10 

1. Studies on Propellants (Final Report), Bryce L. 

Crawford, Jr., OSRD 6374, OEMsr-716, University 
of Minnesota, Oct. 31, 1945. Div. 3-300-MI 

la. Ibid., Volume II, Appendix UM/11. 

lb. Ibid., Volume II, Appendix UM/17. 

lc. Ibid., Volume I, Part II, Section 9. 

l d. Ibid., Volume II, Appendix UM/16. 

2. Interim Ballistic Studies (Final Report), Sidney 
Golden, OSRD 5773, OEMsr-273, ABL Report VI- 
SA, George Washington University, December 1945. 

Div. 3-220-M5 

3. Burning Rate Studies of Double-Base Powder 

(Final Report), William H. Avery, Roy E. Hunt, 
and M. N. Donin, OSRD 5827, OEMsr-273, ABL 
Report P-1, George Washington University, Janu- 
ary 1946. Div. 3-361.21-M6 

4. Erosive Burning of Double-Base Powders (Final 
Report), R. J. Thompson and F. T. McClure, OSRD 
5831, OEMsr-273, ABL Report P-1.1, George Wash- 
ington University, December 1945. 

Div. 3-361.21-M5 



350 


BIBLIOGRAPHY 


5. Effects of Pressure and Temperature on the Rate 

of Burning of Double-Base Powders of Different 
Compositions (Final Report), William H. Avery, 
Roy E. Hunt, and L. D. Sachs, OSRD 5824, OEMsr- 
273, ABL Report P-1.4, George Washington Uni- 
versity, March 1946. Div. 3-361.211-M4 

6. Determination of Burning Rates of Certain Pow- 

ders by the Strand Technique (Final Report), J. J. 
Donovan, L. F. Gonyea, and H. Fritz, OSRD 5833, 
OEMsr-273, ABL Report P-1.2, Service Projects 
NO-33, OD-14, and P-32, George Washington Uni- 
versity, June 1946. Div. 3-355-M7 

7. Studies of Radiation Phenomena in Rockets (Final 

Report), John Beek, Jr., William H. Avery, M. J. 
Dresher, F. T. McClure, and S. S. Penner, OSRD 
5817, OEMsr-273, ABL Report P-2, Service Proj- 
ects OD-14 and NO-33, George Washington Uni- 
versity, June 1946. Div. 3-241-M2 

8. Flame Temperature and Radiation Studies in 

Rockets (Final Report), Ray S. Craig, OSRD 5832, 
OEMsr-273, ABL Report P-2.1, Service Projects 
NO-33 and P-22, George Washington University, 
December 1945. Div. 3-241-MI 

9. The Jet-Assisted Take-Off Unit (Final Report), 

Lyman G. Bonner and William H. Avery, OSRD 
5815, OEMsr-273, ABL Report W-19, Service Proj- 
ects NA-197 and W-191, George Washington Uni- 
versity, December 1945. Div. 3-841-M2 

10. Determination of Burning Rates from Pressure- 

Time Relations in Closed Chambers (Final Report), 
Lyman G. Bonner, OSRD 5816, OEMsr-273, ABL 
Report P-1.3, Service Projects OD-14, NO-33, and 
P-31, George Washington University, December 
1945. Div. 3-355-M5 

11. Propellant Charge Development for 4.5-in. Spinner 

Rockets, T38E5, T105, and T110 (Final Report), 
D. M. Brasted and S. D. Brandwein, OSRD 5800, 
OEMsr-273, ABL Report W-ll, Service Project 
OD-166, George Washington University, December 
1945. Div. 3-320-M6 

12. Bourdon System for Pressure Measurement (Final 

Report), Roy E. Hunt and William H. Avery, 
OSRD 5860, OEMsr-273, ABL Report J-3.1, Serv- 
ice Projects OD-14 and NO-33, George Washington 
University, December 1945. Div. 3-611-M6 

13. Static Range Operational and Fire Control Equip- 
ment for Rocket Research (Final Report), C. M. 
Lathrop and N. E. Alexander, OSRD 5855, OEMsr- 
273, ABL Report J-l, Service Projects OD-14 and 
NO-33, George Washington University, June 1946. 

Div. 3-626-M6 

14. Development of Propellant Charge for 115-mm Air- 
craft Rocket (Final Report), John Beek, Jr., Ray- 
mond L. Arnett, G. W. Engstrom, M. Goldman, and 


A. Kossiakoff, OSRD 5784, OEMsr-273, ABL Re- 
port W-8.1, Service Projects W-80, OD-161, and 
NO-245, George Washington University, June, 
1946. Div. 3-320-M8 

15. Effect of Pressure and Temperature on the Rate of 
Burning of Double-Base Powders of Different Com- 
positions, William H. Avery and Roy E. Hunt, 
OSRD 1993, NDRC Armor and Ordnance Report 
A-225, Section H, Division 3, Service Projects OD- 
14, OD-26, and others, October 1943. 

Div. 3-361. 211-MI 

16. Improvement of Components for l*.5-in. Rocket M8 
(Final Report), D. W. Osborne and B. Weissmann, 
OSRD 5777, OEMsr-273, ABL Report W-4, Serv- 
ice Projects W-40, NO-248, and others, George 
Washington University, December 1945. 

Div. 3-531.3-M2 

17. Development of TU Powder Charge for the 2.36-in. 

Rocket Grenade (Final Report issued by Division 
8), Rufus Lumry and L. N. Streff, OSRD 5589, 
OEMsr-273, ABL Report W-3.4, Service Projects 
OD-14 and OD-200, George Washington University, 
Nov. 29, 1945. Div. 3-320-M4 

18. Ballistic Characteristics and Rocket Design Data 

for Extruded Composite Propellants (Final Re- 
port), Rufus Lumry and L. N. Streff, OSRD 5624, 
OEMsr-273, ABL Report P-10.1, Service Project 
OD-14, George Washington University, December 
1945. Div. 3-370-M3 

19. Theories of the Burning of Colloidal Propellants, 
J. Corner, ARD Theoretical Report 2/43 (OSRD 
Liaison Office Ref. No. WA-1297-5), October 1943. 

20. The Theory of the Burning of Double-Base Rocket 
Powders, Oscar K. Rice, OSRD 5224, NDRC Re- 
port, Division 8, Service Project OD-14, University 
of North Carolina, June 25, 1945. Div. 8-601-M5 

21. The Theory of the Burning of Rocket Powders, 
Oscar K. Rice, OSRD 5774, NDRC Report, Division 
8, Service Project OD-14, Nov. 1, 1945. 

Div. 8-601-M6 

22. The Reduced Specific Impulse of Ideal Gases (Final 

Report), Nancy Marmer and F. T. McClure, OSRD 
5828, OEMsr-273, ABL Report P-3, Service Proj- 
ects OD-14, NO-33, and W-0, George Washington 
University, January 1946. Div. 3-244-M3 

23. Rocket Fundamentals, OSRD 3992, OEMsr-273, 

ABL Special Report 4, George Washington Univer- 
sity, 1944. Div. 3-210-M3 

23a. Ibid., Chap. 3, p. 44. 

24. Flight Ballistics Involved in the Use of Rocket- 
Towed Devices (Final Report), Walter J. Harring- 
ton, OSRD 5888, OEMsr-273, ABL Report B-2.4, 
George Washington University, October 1946. 

Div. 3-592-M2 


BIBLIOGRAPHY 


351 


Chapter 11 

1. High-Explosive Anti-Tank 2.36-in. Rocket (Ba- 

zooka) (Final Report), C. N. Hickman and Sidney 
Golden, OSRD 5771, OEMsr-273, ABL Report W-3, 
Service Project OD-26, George Washington Univer- 
sity, December 1945. Div. 3-551. 1-MI 

2. Interim Ballistic Studies (Final Report), Sidney 

Golden, OSRD 5773, OEMsr-273, ABL Report 
W-3.1, Service Project OD-26, George Washington 
University, December 1945. Div. 3-220-M5 

3. The Development of the T-12 Grenade (Final Re- 

port), D. M. Brasted, OSRD 5776, OEMsr-273, 
ABL Report W-3. 3, George Washington University, 
December 1945. Div. 3-551. 1-M2 

4. Development of T-U Powder Charge for the 2.36-in. 
Rocket Grenade (Final Report issued by Division 
8), Rufus Lumry and L. N. Streff, OSRD 5589, 
OEMsr-273, ABL Report W-3.4, George Washing- 
ton University, Nov. 29, 1945. Div. 3-320-M4 

5. X-Ray Photography of Burning Rocket Propellants 
(Final Report), W. P. Spaulding, Ray S. Craig, and 
Sidney Golden, OSRD 5849, OEMsr-273, ABL Re- 
port J-5, Service Projects W-60.1 and W-80, George 
Washington University, December 1945. 

Div. 3-624. 22-MI 

6. T-59 High-Velocity Rocket Grenade (Final Re- 
port), Sidney Golden, W. P. Spaulding, and L. E. 
Morey, OSRD 5779, OEMsr-273, ABL Report W-6, 
Service Projects W-60, NO-247, and OD-163, 
George Washington University, December 1945. 

Div. 3-551.3-M3 

7. Improvement of Components for h. 5-in. Rocket M8 
(Final Report), D. W. Osborne and B. Weissmann, 
OSRD 5777, OEMsr-273, ABL Report W-4, George 
Washington University, December 1945. 

Div. 3-531. 3-M2 

8. 115-mm Aircraft Rocket (Final Report), R. E. 

Gibson and A. Kossiakoff, OSRD 5781, OEMsr-273, 
ABL Report W-8, George Washington University, 
June 1946. Div. 3-531.3-M3 

9. Development of Propellant Charge for 115-mm 

Aircraft Rocket (Final Report), John Beek, Jr., 
Raymond L. Arnett, G. W. Engstrom, M. Goldman, 
and A. Kossiakoff, OSRD 5784, OEMsr-273, ABL 
Report W-8.1, Service Projects W-80, OD-161, and 
NO-245, George Washington University, June 
1946. Div. 3-320-M8 

10. The Development of a High-Performance Compos- 
ite-Propellant Charge for the 115-mm Aircraft 
Rocket (Final Report), Rufus Lumry and L. N. 
Streff, OSRD 5788, OEMsr-273, ABL Report W-8.4, 
George Washington University, December 1945. 

Div. 3-320-M5 


11. Propellant Charge Development for Jj.5-in. Spinner 

Rockets, T38E5, T105, and T110 (Final Report), 
D. M. Brasted and S. D. Brandwein, OSRD 5800, 
OEMsr-273, ABL Report W-ll, Service Project 
OD-166, George Washington University, Decem- 
ber 1945. Div. 3-320-M6 

12. The Rocket for the Anti-Personnel Mine Clearing 
Snake, Ml (Final Report), C. A. Boyd and R. H. 
Bond, OSRD 5795, OEMsr-273, ABL Report 
W-13.1, Service Projects W-131 and OD-186, George 
Washington University, December 1945. 

Div. 3-592-M4 

13. Investigations of the Use of Rockets to Dispense 

Mine Clearing Hose (Final Report), S. D. Brand- 
wein, C. A. Boyd, and Walter J. Harrington, OSRD 
5796, OEMsr-273, ABL Report W-13.2, Service 
Projects W-132 and OD-186, George Washington 
University, December 1945. Div. 3-592-M5 

14. The Rocket for the Projected Line Charge (Final 
Report), C. A. Boyd, D. Leenov, and Walter J. 
Harrington, OSRD 5799, OEMsr-273, ABL Report 
W-13.3, Service Projects W-136 and OD-186, George 
Washington University, December 1945. 

Div. 3-592-M6 

15. Rocket for Projecting Detonating Cable (Final Re- 
port), C. A. Boyd, Walter J. Harrington, and D. 
Leenov, OSRD 5798, OEMsr-273, ABL Report 
W-13.4, Service Projects W-134 and OD-186, George 
Washington University, January 1946. 

Div. 3-592-M7 

16. The Rocket for Towing Bangalore Torpedoes 
(Final Report), R. H. Bond and C. A. Boyd, OSRD 
5801, OEMsr-273, ABL Report W-13.5, George 
Washington University, November 1945. 

Div. 3-592-M3 

17. The Restriction of Powder Burning (Final Re- 
port) , Amos Turk, Lyman G. Bonner, A. J. Mad- 
den, J. J. Donovan, and William H. Avery, OSRD 
5834, OEMsr-273, ABL Report P-4, December 
1945, George Washington University. 

Div. 3-355-M6 

18. Miscellaneous Propellant Studies (Final Report), 
Lyman G. Bonner, Sidney Golden, and W. P. 
Spaulding, OSRD 5852, OEMsr-273, ABL Report 
P-10, Service Projects NO-33, OD-14, and others, 
George Washington University, December 1945. 

Div. 3-362-M4 

19. The Jet-Assisted Take-Off Unit (Final Report), 

Lyman G. Bonner and William H. Avery, OSRD 
5815, OEMsr-273, ABL Report W-19, Service Proj- 
ects NA-197 and W-191, George Washington Uni- 
versity, December 1945. Div. 3-841-M2 

20. The One-Shot Portable Flame Thrower, E16R1 
(Final Report), Roy E. Hunt, A. Stefcik, L. F. 
Gonyea, and William H. Avery, OSRD 5805, 




352 


BIBLIOGRAPHY 


OEMsr-273, ABL Report W-16.2, Service Project 
CWS-10, George Washington University, Febru- 
ary 1946. Div. 3-821-M5 

21. The Bumblebee Rocket Motor (Final Report), S. S. 
Penner, OSRD 5821, OEMsr-273, ABL Report 
W-22, Service Projects NO-296 and W-221, George 
Washington University, December 1945. 

Div. 3-415-M9 

22. Design of the High-Velocity Rocket (VICAR) 
(Final Report), R. J. Thompson and R. R. Newton, 
OSRD 5793, OEMsr-273, ABL Report W-21, Serv- 
ice Projects W-210 and OD-201, George Washing- 
ton University, December 1945. Div. 3-551. 5-MI 

23. Small-Caliber High-Velocity Rocket (CURATE ) 
(Final Report), R. J. Thompson, G. D. Brewer, and 
R. R. Newton, OSRD 5820, OEMsr-273, ABL 
Report W-21.1, Service Projects OD-201 and W-210, 
George Washington University, January 1946. 

Div. 3-551.4-MI 

24. Studies of the Mechanism of Burning of Double- 
Base Rocket Propellants (Final Report), Farring- 
ton Daniels and associates, OSRD 6559, NDRC 
A-485, OEMsr-762, Service Projects OD-14 and 
NO-33, University of Wisconsin, January 1945. 

Div. 3-361.21-M4 

25. Processing of Rocket Propellants (Final Report), 

W. H. Corcoran and Quentin Elliott, OSRD 2552, 
OEMsr-418, California Institute of Technology, 
1946. Div. 3-300-M2 

26. Dry Extrusion of Powder at Allegany Ballistics 
Laboratory (Final Report), G. F. Padgett and 
Howard E. Higbie, OSRD 5844, OEMsr-273, ABL 
Report P-7, Service Projects OD-14, NO-33, and 
P-70, George Washington University, December 

1945. Div. 3-361. 41-MI 

27. The Drag of the Propellant Gases on the Powder 

Charge in Rockets (Final Report), F. T. McClure, 
J. Barkley Rosser, and James F. Kincaid, OSRD 
5872, OEMsr-273, ABL Report B-1.2, Service Proj- 
ects OD-14 and NO-33, George Washington Uni- 
versity, February 1946. Div. 3-351-MI 

28. Physical Properties of Propellants (Final Report), 
Howard Higbie, OSRD 5845, OEMsr-273, ABL 
Report P-8, George Washington University, De- 
cember 1945. 

Chapter 12 

1. Burning Rate Studies of Double-Base Powder 
(Final Report), William H. Avery, Roy E. Hunt, 
and M. N. Donin, OSRD 5827, OEMsr-273, ABL Re- 
port P-1, Service Projects OD-14, P-10.1, and 
others, George Washington University, January 

1946. Div. 3-361.21-M6 


2. Determination of Burning Rates of Certain Pow- 

ders by the Strand Technique (Final Report), 
J. J. Donovan, L. F. Gonyea, and H. Fritz, OSRD 
5833, OEMsr-273, ABL Report P-1.2, Service Proj- 
ects NO-33, OD-14, and P-32, George Washington 
University, June 1946. Div. 3-355-M7 

3. Determination of Burning Rates from Pressure- 

Time Relations in Closed Chambers (Final Report), 
Lyman G. Bonner, OSRD 5816, OEMsr-273, ABL 
Report P-1.3, Service Projects OD-14, NO-33, and 
P-31, George Washington University, December 
1945. Div. 3-355-M5 

4. Effect of Pressure and Temperature on the Rate 

of Burning of Double-Base Powders of Different 

Compositions (Final Report), William H. Avery, 
Roy E. Hunt, and L. D. Sachs, OSRD 5824, OEMsr- 
273, ABL Report P-1.4, George Washington Uni- 
versity, Mar. 3, 1946. Div. 3-361.211-M4 

5. Rocket Fundamentals , OSRD 3992, OEMsr-273, 

ABL Special Report 4, George Washington Univer- 
sity, 1944. Div. 3-210-M3 

6. Erosive Burning of Double-Base Powders (Final 
Report) , R. J. Thompson and F. T. McClure, OSRD 
5831, OEMsr-273, ABL Report P-1.1, George Wash- 
ington University, December 1945. 

Div. 3-361.21-M5 

7. Design of the High-Velocity Rocket (VICAR) 
(Final Report), R. J. Thompson and R. R. Newton, 
OSRD 5793, OEMsr-273, ABL Report W-21, Serv- 
ice Projects W-210 and OD-201, George Washing- 
ton University, December 1945. Div. 3-551.5-MI 

8. Small-Caliber High-Velocity Rocket (CURATE) 
(Final Report), R. J. Thompson, G. D. Brewer, and 
R. R. Newton, OSRD 5820, OEMsr-273, ABL Re- 
port W-21.1, Service Projects OD-201 and W-210, 
George Washington University, January 1946. 

Div. 3-551.4-MI 

9. Studies of Radiation Phenomena in Rockets (Final 

Report), John Beek, Jr., William H. Avery, M. J. 
Dresher, F. T. McClure, and S. S. Penner, OSRD 
5817, OEMsr-273, ABL Report P-2, Service Proj- 
ect OD-14 and NO-33, George Washington Uni- 
versity, June 1946. Div. 3-241-M2 

10. Flame Temperature and Radiation Studies in 

Rockets (Final Report), Ray S. Craig, OSRD 5832, 
OEMsr-273, ABL Report P-2.1, Service Projects 
OD-14, NO-33, and P-22, George Washington Uni- 
versity, December 1945. Div. 3-241-MI 

11. The Drag of the Propellant Gases on the Powder 
Charge in Rockets (Final Report), F. T. McClure, 
J. Barkley Rosser, and James F. Kincaid, OSRD 
5872, OEMsr-273, ABL Report B-1.2, George 
Washington University, February 1946. 

Div. 3-351-MI 


BIBLIOGRAPHY 


353 


12. Some Problems of Heat Transfer in Rockets (Final 
Report), John Beek, Jr., J. Barkley Rosser, and 
Harry Siller, OSRD 5886, OEMsr-273, ABL Re- 
port B-3, Service Projects OD-26, NO-33, and 
W-6.1, George Washington University, May 1946. 

Div. 3-241.1-M2 

13. Temperature Transients in Walls of Rocket Cham- 
bers (Final Report), E. A. Cook and E. H. 
deButts, Jr., OSRD 5887, OEMsr-273, ABL Report 
B-3.1, George Washington University. 

14. T-59 High-Velocity Rocket Grenade (Final Re- 
port), Sidney Golden, W. P. Spaulding, and L. E. 
Morey, OSRD 5779, OEMsr-273, ABL Report W-6, 
Service Projects W-60, NO-247, and OD-163, 
George Washington University, December 1945. 

Div. 3-551.3-M3 

15. Formulation of Manufacturing Specifications for 
Solid Propellants (Final Report), Raymond L. 
Arnett, OSRD 5851, OEMsr-273, ABL Report P-9, 
Service Projects NO-33, OD-14, and P-90, George 
Washington University, November 1945. 

Div. 3-362-M3 

Chapter 13 

1. Propellant Charge Design of Solid Fuel Rockets 
(Final Report), William H. Avery and John Beek, 
Jr., OSRD 5890, OEMsr-273, ABL Report B-4, 
Service Projects OD-14 and NO-33, George Wash- 
ington University, June 1946. Div. 3-320-M7 


PART IV 
Chapter 14 

1. Considerations Involved in the Design of Short- 

Burning, Long-Range Rockets, Leverett Davis, Jr., 
OEMsr-418, Report IPC 80, California Institute of 
Technology, Sept. 10, 1945. Div. 3-230-MI 

2. Some Operational and Logistical Problems in the 

Use of Rockets, William A. Fowler, OSRD 2366, 
OEMsr-418, Report JNC 16, California Institute of 
Technology, Feb. 1, 1945. Div. 3-220-M4 

2a. Ibid., p. 4. 

3. The Dependence of the Masses of Rocket Com- 

ponents on Their Dimensions, Leverett Davis, Jr., 
OSRD 1260, NDRC A-149, OEMsr-418, Report TAC 
1, California Institute of Technology, Dec. 10, 

1942. Div. 3-414-MI 

4. A Note on the Reasons Why the Same Spin-Stabi- 

lized Rocket Cannot be Used for Very Accurate 
Fire with a Flat Trajectory and for Barrage Pur- 
poses, Leverett Davis, Jr., OEMsr-418, Report OPC 
33, California Institute of Technology, Apr. 4, 
1945. Div. 3-562-M2 


5. Comparison of Fin and Rotational Stabilization of 
Rockets, Charles C. Lauritsen, OEMsr-418, Report 
JPC 15, Service Projects OD-26, NO-33, and others, 
California Institute of Technology, Jan. 25, 1944. 

Div. 3-561-MI 

Chapter 15 

1. Geometric Malalignment in Rockets: Methods of 

Measurement and Correction; Correlation with Ex- 
perimental Results, Thomas L. Lauritsen, L. A. 
Richards, Sylvan Rubin, and J. G. Waugh, OEMsr- 
418, Report JGC 6, California Institute of Tech- 
nology, Aug. 16, 1943. Div. 3-242-M2 

2. Field Testing of Rockets: Range Operations and 
Metric Photography (Final Report), W. N. Arn- 
quist, R. H. Cox, and others, OSRD 2547, OEMsr- 
418, California Institute of Technology, 1946. 

Div. 3-610-M7 

3. Water Entry and Underwater Ballistics of Pro- 
jectiles (Final Report), OSRD 2551, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-713-M5 

4. Principles of Rocket Design, William A. Fowler and 
Thomas L. Lauritsen, monograph (unclassified), 
Chap. 8, “Underwater Behavior of Rockets,” by 
I. S. Bowen. 

Chapter 16 

1. Rocket Fuzes (Final Report), R. B. King, V. K. 
Rasmussen, and others, OSRD 2545, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-430-M5 

la. Ibid., Sec. 5.04 ff. 

lb. Ibid., Chap. 6. 

2. A Point-Initiating Base-Detonating Electromag- 

netic Fuze (Final Report), Allegany Ballistics 
Laboratory, Bell Telephone Laboratories, and Ex- 
plosives Research Laboratory, OSRD 5881, OEMsr- 
273, ABL Report W-6.1, George Washington Uni- 
versity, March 1946. Div. 3-430-M6 

3. Development of Heads and Fuzes for 115-mm Air- 

craft Rocket (Final Report), M. J. Walker, A. 
Kossiakoff, and F. T. McClure, OSRD 5786, 
OEMsr-273, ABL Report W-8.3, George Washing- 
ton University, June 1946. Div. 3-430-M7 

4. Weekly Progress Report, OEMsr-418, Report PMC 
2.49, Part 2, California Institute of Technology, 
1945, pp. 1-6. Div. 3-530-Ml 

5. Weekly Progress Report, OEMsr-418, Report PMC 

2.53, Part 2, California Institute of Technology, 
1945, pp. 4-10. Div. 3-530-Ml 

5a. Ibid., pp. 11-14. 




354 


BIBLIOGRAPHY 


6. Land Service Use of 11. 7 5-in. Aircraft Rockets 

Against Caves , OSRD 2516, OEMsr-418, Report 
JBC 32, California Institute of Technology, Aug. 
15, 1945. Div. 3-531.4-M3 

7. Catalogue: CIT Rocket Fuzes , CIT Fuze Group, 

OEMsr-418, Report UIC 3, California Institute of 
Technology, Jan. 15, 1944. Div. 3-430-M2 

8. Special Fuzes for Rockets, Projector Charges, and 

Miscellaneous Munitions, U. S. Navy Ordnance 
Pamphlet 1017, U. S. Navy Department, June 13, 
1944. Div. 3-430-M3 

9. Tests of Type HIR Fuzes, Bruce H. Rule and 
W. P. Huntley, OEMsr-418, Report IIC 3, Califor- 
nia Institute of Technology, Nov. 9, 1942. 

10. Tests of Type HIR Fuze Modified to Increase Fir- 
ing Sensitivity , Bruce H. Rule and W. P. Huntley, 
OEMsr-418, Report IIC 4, California Institute of 
Technology, Oct. 12 to Dec. 1, 1942. 

11. The Mark 140 Fuze (HIR 3): Tests of Arming 
Depth, Premature Firing, and Sensitivity, Fuze 
Group, OEMsr-418, Report IIC 18, California In- 
stitute of Technology, Sept. 1, 1943. Div. 3-431-MI 

12. Underwater Performance Tests of BuOrd Mk 6 

and BuOrd Mk 8 Projector Charges (Hedgehog) 
with Mk HO Fuze and Protective Cap. Morris Dam 
Report No. 100, R. L. Noland, OEMsr-418, Re- 
port IOC 23, California Institute of Technology, 
Feb. 28, 1944. Div. 3-731.2-M2 

13. Underwater Performance of 7.2-in.-Diameter Fast- 

Sinking Depth Charge with Mark 140 Fuze with 
and without Protective Cap. Morris Dam Report 
No. 97, Bruce H. Rule and W. P. Huntley, OEMsr- 
418, Report IOC 26, California Institute of Tech- 
nology, May 13, 1944. Div. 3-731.3-M2 

14. Underwater Performance of the 7.2-in.-Diameter 
Fast-Sinking Depth Charge with Case Length In- 
creased 1 in. and 3 in. with Mk HO Fuze with and 
without Protective Cap, Bruce H. Rule and W. P. 
Huntley, OSRD 2223, OEMsr-418, Report IOC 26.2, 
California Institute of Technology, Sept. 7, 1944. 

Div. 3-731.3-M4 

15. Underwater Performance of 6-in.-Diameter Mark 

12 Fast-Sinking Depth Charge with Tails of 
Various Sizes and with Mark 140 Fuze and Pro- 
tective Cap, Bruce H. Rule and W. P. Huntley, 
OEMsr-418, Report IOC 27, California Institute of 
Technology, June 7, 1944. Div. 3-731.3-M3 

16. Tests of Preliminary Firing Mechanism for SIR 

Fuze, Bruce H. Rule and W. P. Huntley, OEMsr- 
418, Report IIC 8, California Institute of Technol- 
ogy. Div. 3-433-M3 


17. SIR Fuze Tests, 8 Dec 42 to 15 Jan 43, N. Gunder- 
son, D. E. Brink, C. F. Robinson, V. Rasmussen, 
and R. B. King, OEMsr-418, Report IIC 9, Cali- 
fornia Institute of Technology, January 1943. 

Div. 3-433-MI 

18. Tests of SIR (Mark 139) Fuze. Feb. 27 to April 
27, 1943, OEMsr-418, Report IIC 14, California 
Institute of Technology, April 1943. Div. 3-433-M2 

19. Mk 146 Fuze (PIR) : Static Firing Progress Re- 

port, D. E. Brink (Fuze Group), OEMsr-418, Re- 
port IIC 21, California Institute of Technology, 
Mar. 18, 1944. Div. 3-432-M2 

20. Delay Ejector Units for the 3.5-in. “Window” 
Rocket, Bruce H. Sage, OSRD 2542, OEMsr-418, 
Report JKC 1, California Institute of Technology, 
July 19, 1945. 

21. Jet Propelled Illuminating Flare, W. E. Jeremiah, 

OSRD 992, OEMsr-273 and 256, Service Projects 
OD-26, NO-120, and PA-369, George Washington 
University and Bell Telephone Laboratories, Inc., 
Nov. 12, 1942. Div. 3-810-M2 

22. Thermal Ignition and Arming Elements for Use 
with Rockets, C. N. Hickman, OSRD 1022, NDRC 
A-58 M, Service Projects OD-26, PA-361, and 
others, George Washington University and Bell 
Telephone Laboratories, Inc., Nov. 14, 1942. 

Div. 3-223-MI 


Chapter 17 

1. Rocket Launchers for Surface Use (Final Report), 
Paul E. Lloyd, OSRD 2548, OEMsr-418, California 
Institute of Technology, 1946. Div. 3-491. 2-MI 

2. Firing of Rockets from Aircraft: Launchers, 

Sights, and Flight Tests (Final Report), R. V. 
Adams, C. D. Anderson, and others, OSRD 2549, 
OEMsr-418, California Institute of Technology, 
1946. Div. 3-530-M4 

3. Ripple Firing Mechanism for Launching Rockets 

(Final Report), D. D. Miller and T. H. Guettich, 
OSRD 6158, OEMsr-256, Service Project OD-26, 
BTL, Feb. 9, 1946. Div. 3-491.214-MI 

4. Retro-Bombing : A Description of Projectiles and 

Installations on Aircraft, OEMsr-418, Report JBC 
18, California Institute of Technology, June 23, 
1943. Div. 3-532. 1-MI 

5. Officers’ Manual, Vertical Bombing from PBY-5 

Aircraft, Vertical Bombing Section, UP Group, 
OEMsr-418, Report JNC 4, California Institute of 
Technology, Jan. 19, 1943. Div. 3-532.2-M8 

6. Officers’ Manual: Vertical Bombing from TBF-1 

and TBF-2 Aircraft Using 300-ft/sec Ammunition, 
OEMsr-418, Report JNC 5, California Institute of 
Technology, June 5, 1943. Div. 3-532. 2-M12 


BIBLIOGRAPHY 


355 


7. Weekly Progress Repoi'ts, OEMsr-418, PMC, Cali- 

fornia Institute of Technology, October 1942 to 
May 1943. Div. 3-110-M5 

8. Local Progress Reports , OEMsr-418, NMC, Cali- 
fornia Institute of Technology, October 1942 to 
May 1943. 

9. Effects of Rocket Blast on Aircraft Structures , 

E. C. Briggs and C. H. Wilts, OSRD 2284, OEMsr- 
418, Report JTC 1, California Institute of Tech- 
nology, Nov. 16, 1944. Div. 3-247-MI 

Chapter 18 

1. Ballistic Data , Fin-Stabilized and Spin-Stabilized 
Rockets (Final Report), OSRD 2544, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-220-M6 

2. Rocket Launchers for Surface Use (Final Report), 
Paul E. Lloyd, OSRD 2548, OEMsr-418, California 
Institute of Technology, 1946. Div. 3-491.2-MI 

3. Firing of Rockets from Aircraft: Launchers, 
Sights, Flight Tests (Final Report), R. V. Adams, 
C. D. Anderson, and others, OSRD 2549, OEMsr- 
418, California Institute of Technology, 1946. 

Div. 3-530-M4 

4. Entry and Underwater Characteristics of 7.2-in.- 
Diameter, Flat-Nosed Mousetrap and Hedgehog 
Projectile, Bruce H. Rule, OEMsr-329, Report IBC 
8, California Institute of Technology, July 13, 1942. 

Div. 3-731.1-M4 

5. Effect of Tail Spin on Underwater Projectiles, 
Bruce H. Rule, OEMsr-329, Report IPC 9, Cali- 
fornia Institute of Technology, Aug. 5, 1942. 

Div. 3-714-MI 

6. The Antisubmarine Rocket Projectile and Projec- 
tor, Thomas L. Lauritsen and William R. Smythe, 
NDRC Report A-50, Apr. 27, 1942. Div. 3-492-M3 

7. The Antisubmarine Bomb (ASB), Parts I and II, 
W. N. Arnquist and others, Part I, OSRD 758, 
NDRC A-74I, and Part II, OSRD 803, NDRC 
A-77II, OEMsr-418, Report JBC 7, California In- 
stitute of Technology, June 25, 1942. 

Div. 3-731.1-M3 

8. Use of Mousetrap Ammunition, Thomas L. Laurit- 
sen, OEMsr-418, Report JBC 8, California Insti- 
tute of Technology, June 27, 1942. Div. 3-854-MI 

9. Mousetrap Operating Instructions, L. B. Slicter and 
Thomas L. Lauritsen, OEMsr-418, Report JBC 13, 
California Institute of Technology, Oct. 25, 1942. 

Div. 3-731.1-M8 

10. Elements in the Effectiveness of Antisubmarine 
Attacks by Surface Craft, Norman A. Haskell, et al., 
OSRD 2189, Service Project NO-36.5, California 
Institute of Technology, May 1944. Div. 3-732-M2 


11. Impact and Deceleration of the ASB ( Mousetrap ), 
Bruce H. Rule and W. P. Huntley, IPC 12, Cali- 
fornia Institute of Technology, January 1943. 

Div. 3-731. 1-M13 

12. Water Entity and Underwater Trajectory Tests on 
Bureau of Ordance ASPC (Mousetrap), Bruce H. 
Rule, OEMsr-418, Report IPC 14, California In- 
stitute of Technology, Dec. 3, 1942. 

Div. 3-731. 1-M9 

13. Underwater Performance Tests of 7.2-in. Rocket 
Mark 3 Mousetrap Assembly with Mark 131 Fuze 
and with Mark 140 Fuze and Protective Cap, R. L. 
Noland, OEMsr-418, Report IBC 60, California 
Institute of Technology, Feb. 4, 1944. 

Div. 3-731. 1-M11 

14. Slat Deck Impact Deceleration Tests, ASPC Mk 1, 

Bruce H. Rule and W. P. Huntley, OEMsr-418, 
Report IOC 5, California Institute of Technology, 
July 12, 1943. Div. 3-731.21-M2 

15. Impact and Deceleration of the ASPC Mark 1 
Projectile and Modified AS Bomb, Bruce H. Rule 
and W. P. Huntley, OEMsr-418, Report JBC 15, 
California Institute of Technology, Feb. 10, 1943. 

Div. 3-731.21-MI 

16. Underwater P erf ormance of B.O. ASPC Mark 11, 

Bruce H. Rule and W. P. Huntley, OEMsr-418, 
Report IPC 33, California Institute of Technology, 
Aug. 10, 1943. Div. 3-731.22-MI 

17. Underwater Characteristics of 7.2-in.-Diameter 
Mousetrap Projectile and 7-in.-Diameter Integral 
Motor Modification Tests Mar. 31 to May 15, 1942, 
Bruce H. Rule, OEMsr-418, Report IPC 5, Cali- 
fornia Institute of Technology, no date. 

Div. 3-731.1-MI 

18. AS Projector Charges — Terminal Velocity Sum- 

mary, Bruce H. Rule and W. P. Huntley, OEMsr- 
418, Report IPC 38, California Institute of Tech- 
nology, Aug. 18, 1943. Div. 3-712-MI 

19. Performance Comparisons for 7.2-in.-Diameter 
Modified Mousetrap Projectile, May 28 to June 20, 
1942, Bruce H. Rule, OEMsr-329, Report IPC 7, 
California Institute of Technology, June 1942. 

Div. 3-731. 1-M2 

20. Underwater Tests of Mousetrap with Line and 

Drogue, Bruce H. Rule and W. P. Huntley, OEMsr- 
418, Report IOC 17, California Institute of Tech- 
nology, Aug. 12, 1943. Div. 3-731.1-M10 

21. Development of a Propellant Grain for Use in a 

2-in. Reaction Chamber, Bruce H. Sage, OEMsr- 
418, Report JDC 37, California Institute of Tech- 
nology, Feb. 10, 1943. Div. 3-361.524-M3 

22. The 7.2-in. Rocket Launcher, Mark 20 [and] Mark 

22, and Ammunition, Ordnance Pamphlet 1002, 
Nov. 30, 1943. Div. 3-491.22-MI 


356 


BIBLIOGRAPHY 


23. Manual 7.2-in. Demolition Rocket, OEMsr-418, Re- 
port JBC 24, Service Project OD-137, California 
Institute of Technology, Feb. 10, 1944. 

Div. 3-594-MI 

24. Use of 4-5-in. Barrage Rocket, Thomas L. Lauritsen, 
F. C. Lindvall, and L. A. Richards, OSRD 842, 
NDRC A-85, OEMsr-418, Report JBC 10, Califor- 
nia Institute of Technology, Aug. 1, 1942. 

Div. 3-520-MI 

25. Use of 4.5-in. Barrage Rocket (Revised), UP 
Group, OEMsr-418, Report JBC 10.2, California 
Institute of Technology, Sept. 10, 1942. 

Div. 3-520-M3 

26. Manual: Use of 4-5-in. Barrage Rocket — Second 
Edition, UP Group, OEMsr-418, Report JBC 10.6, 
California Institute of Technology, Apr. 7, 1943. 

Div. 3-520-M9 

27. BR Fragmentation, O. C. Wilson, OEMsr-418, Re- 

port IQC 1, California Institute of Technology, 
July 7, 1943. Div. 3-521-MI 

28. Comparison of Fragmentation of the 4.5-in. Bar- 

rage Rocket with the 105-mm Howitzer Shell, O. C. 
Wilson, C. A. Wirtanen, and J. A. Gilbert, OEMsr- 
418, Report IQC 2, California Institute of Tech- 
nology, July 30, 1943. Div. 3-521-M2 

29. Fragmentation Tests on Special BR Bodies, O. C. 

Wilson, C. A. Wirtanen, and J. A. Gilbert, OEMsr- 
418, Report IQC 3, California Institute of Technol- 
ogy, Aug. 19, 1943. Div. 3-521-M3 

30. Single Shroud Rocket Tail with Internal Insulated 

Firing Ring, L. A. Richards, OEMsr-418, Report 
IBC 11, California Institute of Technology, Jan. 
22, 1943. Div. 3-470-MI 

31. Tests of Lateral Dispersion of BR with Various 
Nozzles, J. G. Waugh and L. A. Richards, OEMsr- 
418, Report IBC 14, California Institute of Tech- 
nology, Nov. 24, 1942 to Jan. 14, 1943. 

Div. 3-520-M7 

32. Tests of Lateral Dispersion of BR with Various 
Nozzles, J. G. Waugh and L. A. Richards, OEMsr- 
418, Report IBC 19, California Institute of Tech- 
nology, Jan. 30 to Feb. 5, 1943. Div. 3-520-M7 

33. Gas-Malalignment and Deflection-Malalignment 

Ratio for All Types of BR Fired from September 
20, 19 42 to April 1, 1943, C. W. Snyder, OEMsr-418, 
Report IBC 23, California Institute of Technology, 
April 1943. Div. 3-520-M8 

34. Experimental Attempts to Improve the Accuracy 

of Rockets, O. C. Wilson and Gabriel E. Kron, 
OEMsr-418, Report JPC 9, California Institute of 
Technology, Nov. 26, 1943. Div. 3-249-M4 


36. “Heavey’s Army,” by Robert Shaplen, Yale Review, 
Spring, 1945. 

37. Projector for the 4 V 2 -in. Barrage Rocket, OEMsr- 

418, Report JEC 4, California Institute of Technol- 
ogy, July 25, 1942. Div. 3-492.2-MI 

38. Manual: Methods of Manufacture for the 4.5-in. 
BR, Lowell Martin, OEMsr-418, Report JSC 1, 
California Institute of Technology, Sept. 22, 1943. 

Div. 3-520-M10 

39. Design of a Fast-Burning Propellant Grain for the 

Barrage Rocket Motor, Bruce H. Sage, OEMsr-418, 
Report JDC 41, California Institute of Technology, 
Mar. 30, 1943. Div. 3-320-MI 

40. BR Parachute Drops, J. E. Thomas and Paul E. 
Lloyd, OEMsr-418, Report IBC 53, California In- 
stitute of Technology, Oct. 20, 1943. 

Div. 3-522-MI 

41. BR Parachute Drops, Paul E. Lloyd and R. D. 
Ridgeway, OEMsr-418, Report IBC 57, California 
Institute of Technology, Dec. 6, 1943. 

Div. 3-522-M2 

42. Weekly Progress Report, OEMsr-418, Report PMC 
1.20, California Institute of Technology, Mar. 15, 

1942. Div. 3-110-M2 

43. Comparison of Design and Performance of 7-in. 

CWR and 7.2-in. CWR-N, Thomas L. Lauritsen, 
OEMsr-418, Report IBC 47, California Institute of 
Technology, Sept. 14, 1943. Div. 3-571-MI 

44. Accuracy of the CWR-N, C. Weinland, J. W. Mc- 

Connell, and F. W. Thiele, OEMsr-418, Report IBC 
39, California Institute of Technology, Sept. 25, 

1943. Div. 3-573-M3 

45. Dispersion and High Temperature Limit of the 
CWR-N, Projectile Section, OEMsr-418, Report 
IBC 62, California Institute of Technology, Feb. 15, 

1944. 

46. The Chemical Warfare Bomb (CWB), R. B. King 
and W. H. Sleeper, OSRD 866, NDRC A-86, 
OEMsr-418, Report JBC 11, California Institute 
of Technology, Aug. 20, 1942. Div. 3-570-M4 

47. Design of a CWR Grain for the 3.25-in. Mk 5 

Motor, Quentin Elliott, OEMsr-418, Report IDC 
34, California Institute of Technology, Aug. 5, 
1943. Div. 3-320-M2 

48. History of Solventless Extrusion of Double-Base 
Propellant at the California Institute of Technol- 
ogy, Bruce H. Sage, OEMsr-418, Report IDC 43, 
California Institute of Technology, Mar. 1, 1945. 

Div. 3-361-MI 

49. Rocket Launchers Mk 17 and Mk 17 Mod 1 , Ord- 
nance Pamphlet 1133, Aug. 30, 1944. 

Div. 3-491.233-MI 


35. The Military Engineer, May 1944 and July 1945. 


BIBLIOGRAPHY 


357 


50. Rocket Targets as of November 1, 1941, A. J. 

Dempster, NDRC Report A-27, OSRD 311, Service 
Project OD-26, Dec. 24, 1941. Div. 3-625-MI 

51. Rocket Targets, William A. Fowler, OSRD 415, 

NDRC A-34, OEMsr-418, Report JBC 3, Research 
Project PDRC-155, California Institute of Tech- 
nology, Jan. 31, 1942. Div. 3-625-M3 

52. Rocket Targets, William A. Fowler, OSRD 415, 

NDRC A-34, OEMsr-418, Report JBC 4, Research 
Project PDRC-155, California Institute of Tech- 
nology, Apr. 14, 1942. Div. 3-625-M3 

53. CIT Rocket Targets, James B. Edson, OEMsr-418, 

Report JBC 17, California Institute of Technology, 
June 12, 1943. Div. 3-625-M6 

54. Feasibility of Visual Coincidence Scoring by Two 
Observers of Tracer Bullets Shot at Rocket Tar- 
gets, Jesse W. DuMond, OEMsr-250, Report JNC 3, 
California Institute of Technology, May 1, 1942. 

Div. 3-625.2-M3 

55. Use of the 7-Dial Scoring Register and Tape Re- 

corder for Rocket Target Practice, James B. Edson, 
OEMsr-418, Report JNC 6, California Institute of 
Technology, Aug. 11, 1943. Div. 3-625.2-M4 

56. Gunnery and Tactical Training with Rocket Tar- 
gets, James B. Edson, OEMsr-418, Report JNC 10, 
California Institute of Technology, Dec. 14, 1943. 

Div. 3-625-M7 

57. Manual: Manufacturing and Inspection Problems: 

Rocket Target Mk 3, Developmental Engineering 
Section, OEMsr-418, Report JSC 3, Service Proj- 
ect NO-170, California Institute of Technology, 
Mar. 1, 1944. Div. 3-625-M8 

58. The Effect of Fin Size, Burning Time, and Projec- 

tor Length on the Accuracy of Rockets, Ira S. 
Bowen, Leverett Davis, Jr., and Leon Blitzer, 
OSRD 1330, NDRC A-164, OEMsr-418, Report 
JPC 3, California Institute of Technology, March 
1943. Div. 3-249-M3 

59. Chemical Warfare Grenade (CWG), Sylvan Rubin, 

OEMsr-418, Report IBC 2, California Institute of 
Technology, Feb. 5, 1942. Div. 3-570-MI 

60. The Chemical Warfare Grenade, R. B. King, Sylvan 

Rubin, and O. C. Wilson, OSRD 585, NDRC A-57, 
OEMsr-250, Report JBC 6, California Institute of 
Technology, May 20, 1942. Div. 3-570-M3 

61. The Theory of the Variation with Temperature of 

the Dispersion of the CWG, Leverett Davis, Jr., 
OEMsr-418, Report MTC 4, California Institute of 
Technology, May 1, 1942. Div. 3-573-MI 

62. Effect of Fins on the Yaw and Deflection of GWG’s, 
Leon Blitzer, OEMsr-418, Report MTC 5, Cali- 
fornia Institute of Technology, June 1, 1942. 

Div. 3-573-M2 


63. Use of Subcaliber Mousetrap Ammunition, O. C. 
Wilson, OEMsr-418, Report JBC 9, California In- 
stitute of Technology, June 30, 1942. 

Div. 3-854-M2 

64. Underwater Tests of Proposed Practice Hedgehog 

Ammunition, Bruce H. Rule and W. P. Huntley, 
OEMsr-418, Report IOC 4, California Institute of 
Technology, June 2, 1943. Div. 3-855-MI 

65. BuOrd Subcaliber Ammunition for AS Projector 

Mark 10 (Hedgehog) Underwater Performance 
Tests, Bruce H. Rule and W. P. Huntley, OEMsr- 
418, Report IOC 11, California Institute of Tech- 
nology, Sept. 21, 1943. Div. 3-731.2-MI 

66. Performance Characteristics of 2 V 2 -in.-Diameter 
Underwater Integral Motor Target Bomb, Bruce 
H. Rule, OEMsr-418, Report IPC 6, California 
Institute of Technology, May 29, 1942. 

Div. 3-723-MI 

67. VAR Subcaliber Underwater Performance Tests, 

Bruce H. Rule and W. P. Huntley, OEMsr-418, Re- 
port IPC 27, California Institute of Technology, 
no date. Div. 3-721-M4 

68. ASPC Subcaliber Underwater Performance Tests, 

Bruce H. Rule and W. P. Huntley, OEMsr-418, 
Report IPC 32, California Institute of Technology, 
no date. Div. 3-731.2-M3 

69. Underwater Tests of 2 V 2 -in. ASR Subcaliber with 
Magnesium Flare Head, Bruce H. Rule and W. P. 
Huntley, OEMsr-418, Report IPC 37, California 
Institute of Technology, Aug. 3, 1943. 

Div. 3-721-MI 

70. Water Tunnel Tests of the 7.2-in. Chemical Rocket, 
Robert T. Knapp and Harold L. Doolittle, Section 
No. 6.1-sr-207-1261, HML Rep. ND-22, California 
Institute of Technology, Dec. 22, 1943. 

Div. 6-722.7-M4 

Chapter 19 

1. Vertical Bombing as of November 25, 1942, UP 
Design and Field Testing Group, NDRC Armor 
and Ordnance Report A-141, California Institute 
of Technology, p. 2, Feb. 5, 1943. Div. 3-532.2-M9 

2. Ballistic Data, Fin-Stabilized and Spin-Stabilized 
Rockets (Final Report), OSRD 2544, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-220-M6 

3. Retro-Bombing : A Description of Projectiles and 

Installations on Aircraft, OEMsr-418, Report JBC 
18, California Institute of Technology, June 23, 
1943. Div. 3-532.1-MI 

4. Weekly Progress Report, OEMsr-418, Report PMC 

2.48, Part 2, California Institute of Technology, 
Sept. 3, 1944, p. 27. Div. 3-110-M9 



358 


BIBLIOGRAPHY 


5. Vertical Antisubmarine Bomb (VASB); Vertical 
Flare (VF), William A. Fowler, OSRD 768, NDRC 
A-76, OEMsr-418, Report OBC 14.2, California In- 
stitute of Technology, July 23, 1942. 

Div. 3-532. 2-MI 

6. Vertical Bombing, C. D. Anderson, W. N. Arnquist, 
and F. C. Lindvall, OSRD 911, NDRC A-54M, 
OEMsr-418, Report OBC 14.4, California Institute 
of Technology, Sept. 12, 1942. Div. 3-532. 2-M5 

7. Vertical Bombing, Third Report, C. D. Anderson 

and others, OSRD 1242, NDRC A-141, OEMsr-418, 
Report OBC 14.5, California Institute of Tech- 
nology, Nov. 25, 1942. Div. 3-532.2-M7 

8. Status Report on Smoke Float Rockets, Sylvan 
Rubin, OEMsr-418, Report IBC 32, California In- 
stitute of Technology, June 7, 1943. Div. 3-810-M3 

9. Smoke Float Rocket, Sylvan Rubin, OEMsr-418, 

Report IBC 54, California Institute of Technology, 
Nov. 24, 1943. Div. 3-810-M4 

10. The 100-Knot Vertical Flare Mark U, John Mc- 
Morris, OEMsr-418, Report JBC 12, California In- 
stitute of Technology, Sept. 21, 1942. 

Div. 3-810-MI 

11. Design of a Cruciform Charge for the 3.25-in. 
Motor, Bruce H. Sage, OEMsr-418, Report JDC 46, 
California Institute of Technology, July 19, 1943. 

Div. 3-322-MI 

12. Development of the Mk 13 Cruciform Propellant 

Grain, Bruce H. Sage, OEMsr-418, Report JDC 
56, California Institute of Technology, Dec. 29, 

1943. Div. 3-322-M2 

13. Weekly Progress Report, OEMsr-418, Report PMC 

2.6, California Institute of Technology, Nov. 14, 

1943, p. 16. Div. 3-110-M5 

14. Brief History of the Development of the 3.5-in. 
Aircraft Rocket, OEMsr-418, Report JBC 25, Cali- 
fornia Institute of Technology, May 10, 1944. 

Div. 3-531.2-M2 

15. Development of the 3.5-in. Aircraft Rocket, Models 
1, 5, and 1U, OSRD 2107, OEMsr-418, Report JBC 
26, California Institute of Technology, June 1, 1944. 

Div. 3-531.2-M3 

16. Underwater Trajectories of 3.5-in. Aircraft Rocket 

Model 5, R. V. Adams, OSRD 2161, OEMsr-418, 
Report JPC 21, California Institute of Technology, 
June 10, 1944. Div. 3-711-M2 

17. Ammunition Manual for the U.5-in. BR (1100-yd) , 

OEMsr-418, Report JBC 19, California Institute of 
Technology, July 26, 1943. Div. 3-852-MI 

18. Further Investigations of the Underwater Be- 
havior of Aircraft Rockets, Ira S. Bowen, R. V. 
Adams, and Sylvan Rubin, OSRD 2152, OEMsr- 


418, Report JBC 27, California Institute of Tech- 
nology, June 26, 1944. Div. 3-721-M3 

19. 3.5-in. AR Bodies with Nonricochet Properties at 

Low Angles of Water Impact, Ira S. Bowen, 
OEMsr-418, Report IPC 64, California Institute of 
Technology, Oct. 10, 1944. Div. 3-713-M2 

20. Manufacturing Methods for 3.25-in. Rocket Motor 

Mk 7 and 3.5-in. Rocket Body Mk 1, CIT Section 
“B,” OEMsr-418, Report JSC 2, Service Project 
NO-170, California Institute of Technology, Jan. 
20, 1944. Div. 3-415-M4 

21. Manual: Inspection Procedures for 3.5-in. Aircraft 

Rocket Model 5 (3.25-in. Rocket Motor Mk 7 and 
3.5-in. Rocket Body Mk 1 ), Developmental Engi- 
neering Section, OSRD 2110, OEMsr-418, Report 
JSC 5, California Institute of Technology, Apr. 29, 
1944. Div. 3-531.2-MI 

22. The Exterior Ballistics of Fm-Stabilized Aircraft 
Rockets, Leon Blitzer and Leverett Davis, Jr., 
OSRD 2528, OEMsr-418, Report JPC 24, Cali- 
fornia Institute of Technology, Aug. 20, 1945. 

Div. 3-561-M2 

23. Dispersion due to Malalignment of Fin-Stabilized 
Rockets in Forward Firing from Aircraft, L. Ivan 
Epstein, OSRD 2190, OEMsr-418, Report JPC 
23, Service Projects OD-16, NO-170, and others, 
California Institute of Technology, Aug. 10, 1944. 

Div. 3-561.1-M3 

24. Trajectories of Aircraft Rockets, 3.5-in. and 5.0-in., 
OSRD 2225, OEMsr-418, Report UBC 27, Califor- 
nia Institute of Technology, Sept. 25, 1944. 

Div. 3-531.2-M4 

25. Airborne Forward-Firing Rockets: Training Notes 
for Aviation Ordnancemen, issued by Aviation 
Training Division, Office of the Chief of Naval 
Operations, U. S. Navy, Conavaer 00-805-48, Opnav 
33-NY-45, May 1945. 

26. Principles of Rocket Firing from Aircraft, Illus- 
trated, OSRD 2428, OEMsr-418, Report JNC 30, 
California Institute of Technology, Apr. 2, 1945. 

Div. 3-530-M3 

27. Aircraft Rockets in Antisubmarine Warfare with 
Special Reference to TBF and PV Aircraft, pre- 
pared by Anti-Submarine Development Detach- 
ment, Air Force, Atlantic Fleet, May 1, 1944. 

28. Forward Firing of 3.5-in. and 5.0-in. Aircraft 

Rockets from TBF-1, PV-1, SBD-5, and F6F-3 Air- 
craft, OEMsr-418, Report JNC 9.3, Service Project 
NO-170, California Institute of Technology, Dec. 
31, 1943. Div. 3-532.31-M2 

29. Handling of Forward-Firing Rocket Equipment 
Aboard Carriers, Commander Fleet Air West 
Coast, OEMsr-418, Report JNC 13, California In- 
stitute of Technology, Jan. 8, 1944. 

Div. 3-532.32-MI 


BIBLIOGRAPHY 


359 


30. CIT Rockets , OEMsr-418, Report UBC 1, California 
Institute of Technology, Nov. 18, 1942. 

Div. 3-510-MI 

31. The 2.25-in. Subcaliber Aircraft Rockets Models 
1 and 3, OSRD 2305, OEMsr-418, Report JBC 30, 
California Institute of Technology, Nov. 20, 1944. 

Div. 3-531. 1-M2 

32. Manual: Inspection Procedures for 2.25-in. Air- 

craft Rocket Model 3 ( Subcaliber ), (2.25-in. Rocket 
Motor Mk 12 and 2.25-in. Rocket Body Mk 1), 
Developmental Engineering Section, OSRD 2119, 
OEMsr-418, Report JSC 6, California Institute of 
Technology, May 30, 1944. Div. 3-531.1-MI 

33. Tests of the 5 " HVAR Projectile with Fin and 
Ring Tails, Harold L. Doolittle, NDRC 2241, CIT 
Hydrodynamics Laboratory, California Institute of 
Technology, Aug. 20, 1945. 

34. Rocket Fuzes (Final Report), R. B. King, V. K. 
Rasmussen, and others, OSRD 2545, OEMsr-418, 
California Institute of Technology. Div. 3-430-M5 

35. “Army Forward-Firing Rocket Launchers” in Con- 
fidential Bulletin, OSRD 2389, OEMsr-418, Report 
LMC 1.16, California Institute of Technology, Feb. 

15, 1945. Div. 3-110-M11 

36. Confidential Bulletin, OSRD 2304, OEMsr-418, Re- 

port LMC 1.12, California Institute of Technology, 
Dec. 1, 1944. Div. 3-110-M10 

37. Development of a 24-lb Cruciform Charge for the 

5.0-in. Rocket Motor, Bruce H. Sage, OSRD 2108, 
OEMsr-418, Report JDC 62, California Institute 
of Technology, May 4, 1944. Div. 3-322-M3 

38. Manual: Description and Use of the 5.0-in. HVAR, 

Models 13A and 14A, OSRD 2291, OEMsr-418, 
Report JBC 29, Service Projects OD-162, OD-164, 
and NO-170, California Institute of Technology, 
Nov. 14, 1944. Div. 3-551.2-M5 

39. Manual: Inspection Procedures for 5.0-in. High- 
Velocity Aircraft Rocket: Models 13, 14, 15, and 

16. (5.0-in. Rocket Motor Mk 1 and 5.0-in. Rocket 
Body Mk 5), Developmental Engineering Section, 
OSRD 2204, OEMsr-418, Report JSC 7, California 
Institute of Technology, Aug. 28, 1944. 

Div. 3-551.2-M2 

40. Manual: Inspection Procedures for 5.0-in. High- 

Velocity Aircraft Rocket, Supplement No. 1. In- 
spection of 5.0-in. Rocket Body Mk 5 Mod 1 , De- 
velopmental Engineering Section, OSRD 2234, 
OEMsr-418, Report JSC 7, Sup. 1, Service Projects 
OD-162 and NO-170, California Institute of Tech- 
nology, September 1944. Div. 3-551.2-M3 

41. Instruction Manual: Optical Inspection Fixture, 
Model M-4, 5.0-in. High-Velocity Aircraft Rocket, 
Developmental Engineering Section, OSRD 2236, 


OEMsr-418, Report JSC 8, Service Projects OD-162 
and NO-170, California Institute of Technology, 
Oct. 2, 1944. Div. 3-551.2-M4 

42. Specifications for Standard Assembly of 5.0-in. 

Rocket Motor CIT Model 38 (5MA5), M. C. Pond, 
OEMsr-418, Report IAC 17, California Institute of 
Technology, May 29, 1945. Div. 3-415-M7 

43. Rocket Launchers for Surface Use (Final Report), 

Paul E. Lloyd, OSRD 2548, OEMsr-418, California 
Institute of Technology, 1946. Div. 3-491.2-MI 

44. Design and Development of the 11.75-in. Rocket 
Motor, C. W. Snyder, OEMsr-418, Report IBC 75, 
California Institute of Technology, Nov. 6, 1945. 

Div. 3-415-M8 

45. 11.75-in. Rocket Ammunition: Description and In- 

structions for Use (Preliminary), Ordnance 
Pamphlet No. 1227, Bureau of Ordnance, July 13, 
1944. Div. 3-853-MI 

46. Land Service Use of 11.75-in. Aircraft Rockets 

Against Caves, OSRD 2516, OEMsr-418, Report 
JBC 32, California Institute of Technology, Aug. 
15, 1945. Div. 3-531.4-M3 

47. Manual: Description and Use of the 11.75-in. Air- 
craft Rocket Model 3 from F4U-1D Aircraft with 
Displacement Launcher, OSRD 2313, OEMsr-418, 
Report JEC 21, Service Project NO-256, California 
Institute of Technology, Dec. 15, 1944. 

Div. 3-491.1-M2 

48. Confidential Bulletin, OSRD 2406, OEMsr-418, 

Report LMC 1.17, California Institute of Technol- 
ogy, Mar. 1, 1945. Div. 3-530-M2 

49. Manual: Inspection Procedures for 11.75-in. Rocket 

Motor Mk 1 , OSRD 2311, OEMsr-418, Report JSC 
9, Service Project NO-256, California Institute of 
Technology, Nov. 22, 1944. Div. 3-415-M6 

Chapter 20 

1. Rocket Fuzes (Final Report), R. B. King, V. K. 
Rasmussen, and others, OSRD 2545, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-430-M5 

la. Ibid., Chap. 6. 

2. Firing of Rockets from Aircraft: Launchers, 

Sights, and Flight Tests (Final Report), R. V. 
Adams, C. D. Anderson, and others, OSRD 2549, 
OEMsr-418, California Institute of Technology, 
1946, Chap. 5. Div. 3-530-M4 

3. Field Testing of Rockets; Range Operations and 

Metric Photography (Final Report), W. N. Arn- 
quist, R. H. Cox, and others, OSRD 2547, OEMsr- 
418, California Institute of Technology, 1946, 
Chap. 4. Div. 3-610-M7 




360 


BIBLIOGRAPHY 


4. Ballistic Data, Fin-Stabilized and Spin-Stabilized 
Rockets (Final Report), OSRD 2544, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-220-M6 


PART Y 

Chapter 21 

1. The Effective Temperatures of Rocket Motors with 

Cruciform Grains, Leverett Davis, Jr., F. E. 
Roach, and J. M. Schmidt, OSRD 2176, OEMsr-418, 
Report JNC 22, California Institute of Technology, 
Aug. 5, 1944. Div. 3-410-M5 

2. Drag Characteristics of Various Aircraft Rocket 
Projectiles, Hsue-Shen Tsien and Leverett Davis, 
Jr., OEMsr-418, Report IPC 52, California Institute 
of Technology, Jan. 15, 1944. Div. 3-243.3-MI 

3. Exterior Ballistic Tables Based on Numerical 
Integrations, prepared by Ordnance Department, 
U. S. Army, Volume 3. 

4. Formulas for the Spin Produced by Inclined Jets, 
Leverett Davis, Jr., OEMsr-418, Report IPC 55, 
California Institute of Technology, Mar. 25, 1944. 

Div. 3-243.2-MI 

5. Qualitative Discussion of Equilibrium Yaw, Sylvan 
Rubin, OEMsr-418, Report OPC 15, California 
Institute of Technology, Dec. 22, 1943. 

Div. 3-243.1-M3 

6. The Effect of Pin Size, Burning Time, and Pro- 
jector Length on the Accuracy of Rockets, Ira S. 
Bowen, Leverett Davis, Jr., and Leon Blitzer, 
OSRD 1330, NDRC A-164, OEMsr-418, Report JPC 
3, California Institute of Technology, Jan. 4, 1943. 

Div. 3-249-M2 

7. Dispersion of Fin-Stabilized Rockets, William A. 
Fowler, OEMsr-418, Report JPC 16, California 
Institute of Technology, Jan. 28, 1944. 

Div. 3-561. 1-MI 

8. Ballistic Data, Fin-Stabilized and Spin-Stabilized 

Rockets (Final Report), OSRD 2544, OEMsr-418, 
California Institute of Technology, 1946, pp. 380- 
383. Div. 3-220-M6 

9. The Effect of Wind on Ground-Fired Spin-Sta- 
bilized Rockets During Burning, James W. Follin, 
Jr., OEMsr-418, Report IPC 76, California Insti- 
tute of Technology, Apr. 23, 1945. Div. 3-243.1-M5 

10. Effect of Wind During Burning, James W. Follin, 
OEMsr-418, Report OPC 15.1, California Institute 
of Technology, Mar. 6, 1945. Div. 3-243. 1-M4 

11. A Note on the Reasons Why the Same Spin- 
Stabilized Rocket Cannot be Used for Very Accu- 
rate Fire with a Flat Trajectory and for Barrage 


Purposes, Leverett Davis, Jr., OEMsr-418, Report 
OPC 33, California Institute of Technology, Apr. 
4, 1945. Div. 3-562-M2 

Chapter 22 

1. The Interior Ballistics of Rockets, R. N. Wimpress, 
monograph (unclassified) to be published by the 
McGraw-Hill Book Company, Inc. 

la. Ibid., Chap. 3. 

lb. Ibid., Chap. 5. 

lc. Ibid., Chap. 7, equation (4). 

l d. Ibid., Chap. 3, Table 1. 

le. Ibid., Chaps. 2, 4, 5, 6, and 7. 

l f. Ibid., Chap. 8. 

lg. Ibid., Chap. 9. 

lh. Ibid., Chap. 12. 

li. Ibid., Chap. 11. 

lj. Ibid., Chap. 10. 

2. Rocket Fundamentals, OSRD 3992, OEMsr-273, 

ABL Special Report 4, George Washington Uni- 
versity, 1944. Div. 3-210-M3 

2a. Ibid., Chap. 2. 

3. Pressure Distribution along Radial-Burning Pro- 
pellant Grains, Bruce H. Sage, OSRD 818, NDRC 
A-84, OEMsr-418, Report JDC 14, California In- 
stitute of Technology, Aug. 10, 1942. 

Div. 3-341.1-MI 

4. The Computation of Pressure-Time Curves, C. T. 
Elvey, OEMsr-418, Report ILC 3, California Insti- 
tute of Technology, June 3, 1943. Div. 3-248-M3 

5. Influence of Size of the Axial Perforation upon 

the Performance of Radial-Burning Grains, Bruce 
H. Sage, OSRD 966, NDRC A-106, OEMsr-418, 
Report JDC 19, California Institute of Technol- 
ogy, Sept. 21, 1942. Div. 3-341.1-M3 

6. Effect of Dimensions on Performance of Tubular 

Grains for 2.25 " Rocket Motors, Bruce H. Sage, 
OEMsr-418, Report JDC 59, Service Project NO- 
33, California Institute of Technology, Jan. 27, 
1944. Div. 3-320-M3 

7. Motor Catalogue Supplementary Internal Ballis- 
tics Curves, OEMsr-418, Report UAC 1.2, Califor- 
nia Institute of Technology, Nov. 20, 1942. 

Div. 3-221-MI 

8. Burning Characteristics in the Axial Perforations 

of Extruded Ballistite Grains, Bruce H. Sage, 
OSRD 815, NDRC A-83, OEMsr-418, Report JDC 
13, California Institute of Technology, Aug. 30, 
1942. Div. 3-361.514-M3 

9. Stabilization of Reaction of Tubular Propellant 
Grains by the Use of Longitudinal Ridges in the 
Central Perforations, Bruce H. Sage, OSRD 2541, 
OEMsr-418, Report JDC 75, California Institute of 
Technology, May 19, 1945. Div. 3-361. 524-M6 


BIBLIOGRAPHY 


361 


10. Available Propellant Shapes, OEMsr-418, Report 

IDC 19, California Institute of Technology, July 
20, 1942. Div. 3-341-MI 

11. Ballistic Data: Fin-Stabilized and Spin-Stabilized 
Rockets (Final Report), OSRD 2544, OEMsr-418, 
California Institute of Technology, 1946. 

Div. 3-220-M6 

12. Tentative Design of a Cruciform Charge for the 

3.25-in. Motor, Bruce H. Sage, OEMsr-418, Re- 
port JDC 46, California Institute of Technology, 
July 19, 1943. Div. 3-322-MI 

13. Development of the Mk 13 Cruciform Propellant 
Grain, Bruce H. Sage, OEMsr-418, Report JDC 56, 
Service Projects OD-26 and NO-33, California 
Institute of Technology, Dec. 29, 1943. 

Div. 3-322-M2 

14. Development of a 24-lb Cruciform Charge for the 

5.0-in. Rocket Motor, Bruce H. Sage, OSRD 2108, 
OEMsr-418, Report JDC 62, California Institute 

of Technology, May 4, 1944. Div. 3-322-M3 

15. Design of Dies for the Extrusion of Solventless 

Ballistite, Bruce H. Sage, OEMsr-418, Report JDC 
44, California Institute of Technology, May 29, 
1943. Div. 3-361. 522-MI 

16. Development of a Triform Grain for 3.25-in. Rocket 
Motors, Bruce H. Sage, OSRD 2535, OEMsr-418, 
Report JDC 79, Service Projects OD-14 and NO-33, 
California Institute of Technology, Sept. 14, 1945. 

Div. 3-321-MI 

17. Some Calculations and Experimental Measure- 
ments Upon the Pressure Distribution Around 
Thin-Webbed Charges During Firing, R. N. Wim- 
press, G. W. Miller, Bruce H. Sage, and William 
N. Lacey, OEMsr-250, Report IDC 10, California 
Institute of Technology, Apr. 8, 1942. 

Div. 3-355-MI 

18. Burning Rate of Four-Spoke Grains of Extruded 

Ballistite, Bruce H. Sage. OEMsr-418, Report 
JDC-18, California Institute of Technology, Sept. 
25, 1942. Div. 3-361.514-M6 

19. The Relation of Column Strength to the Ballistic 

Performance of Mk 13 Grains, Bruce H. Sage, 
OEMsr-418, Report JDC 87, California Institute 
of Technology, Dec. 15, 1945. Div. 3-322-M4 

20. Some Factors Entering into the Design of High- 

Performance Rockets, E. Ellis and F. E. Roach, 
OEMsr-418, Report I AC 5, California Institute of 
Technology, Jan. 10, 1943. Div. 3-415-M3 

21. Some Physical Properties of Double-Base Powders, 
Bruce H. Sage, OEMsr-418, Report JDC 51, Cali- 
fornia Institute of Technology, Oct. 12, 1943. 

Div. 3-361.2-M2 


22. The Investigation of a High-Strength Propellant, 
Bruce H. Sage, OSRD 2364, OEMsr-418, Report 
JDC 67, Service Projects OD-14 and NO-33, Cali- 
fornia Institute of Technology, Nov. 7, 1944. 

Div. 3-362-M2 

23. Investigation of JPH Propellant Lots FDAP 28 

and FDAP 29, Bruce H. Sage, OSRD 2517, OEMsr- 
418, Report JDC 74, Service Projects OD-14 and 
NO-33, California Institute of Technology, June 
5, 1945. Div. 3-330-M3 

24. Compressive, Torsional, and Shear Characteristics 
of Some Double-Base Propellants, Bruce H. Sage, 
OEMsr-418, Report JDC 81, California Institute 
of Technology, Nov. 1, 1945. Div. 3-361. 51-M9 

25. Dependence of the Mass of Propellant in a Rocket 
Motor on the Web Thickness and the Motor Dimen- 
sions, Leverett Davis, Jr., and Chester D. Mills, 
OSRD 1319, NDRC A-163, OEMsr-418, Report JAC 
4, Service Projects OD-26, CWS-22, and NO-33, 
California Institute of Technology, March 1943. 

Div. 3-414-M4 

26. The Use of Ballistite Turnings in Primers (Pre- 

liminary Report), Bruce H. Sage and William N. 
Lacey, OSRD 577, NDRC A-56, OEMsr-250, Re- 
port JCC 2, California Institute of Technology, 
Mar. 14, 1942. Div. 3-361.53-MI 

27. Preliminary Investigation of Metal-Oxidant Ignit- 

ers for Ballistite, Bruce H. Sage, OEMsr-418, Re- 
port JCC 6, California Institute of Technology, 
Feb. 25, 1943. Div. 3-420-M3 

28. Effect of Squib Boosters on the Performance of 

Black Powder Igniters, Bruce H. Sage, OEMsr-418, 
Report JCC 9, California Institute of Technology, 
Aug. 14, 1943. Div. 3-421-MI 

29. Description of an Igniter for Mousetrap Propellant, 

OEMsr-418, Report ICC 1, California Institute of 
Technology, Aug. 10, 1942. Div. 3-420-M2 

30. A Preliminary Investigation of Plastic Cases for 
Igniters for Ballistite, Bruce H. Sage, OSRD 1191, 
NDRC A-138, OEMsr-418, Report JCC 3, Cali- 
fornia Institute of Technology, Sept. 15, 1942. 

Div. 3-422.1-MI 

31. Investigation of the Use of Plastic-Case Igniters 
for the ASPC Motor, Bruce H. Sage, OSRD 1318, 
NDRC A-158, OEMsr-418, Report JCC 5, Califor- 
nia Institute of Technology, Jan. 7, 1943. 

Div. 3-422.1-M2 

32. Development of Cellulose Acetate Igniter Cases for 

1.25-in. and 2.25-in. Rocket Motors, Bruce H. Sage, 
OEMsr-418, Report JCC 8, California Institute of 
Technology, Aug. 12, 1943. Div. 3-422. 1-M5 

33. Performance Tests on Electric Squibs and Rocket 
Igniters After Storage at Elevated Temperatures, 



362 


BIBLIOGRAPHY 


Bruce H. Sage, OEMsr-418, Report JCC 10, Cali- 
fornia Institute of Technology, Oct. 16, 1943. 

Div. 3-421-M2 

34. Threaded-Closure Plastic-Case Igniter for 2.25-in. 

Rocket Motors , Bruce H. Sage, OEMsr-418, Report 
JCC 11, California Institute of Technology, Mar. 
16, 1944. Div. 3-422.1-M6 

35. Development of a Toroid Igniter for Application 
in the 3.25-in. Spin-Stabilized Rocket Motor Mk 13, 
Bruce H. Sage, OEMsr-418, Report ICC 3, Cali- 
fornia Institute of Technology, Nov. 15, 1945. 

Div. 3-420-M5 

36. Design of Box Grids, Sylvan Rubin, OEMsr-418, 

Report IAC 4, California Institute of Technology, 
Nov. 6, 1942. Div. 3-411-MI 

Chapter 23 

1. The Interior Ballistics of Rockets, R. N. Wimpress, 
monograph (unclassified) to be published by the 
McGraw-Hill Book Company, Inc. 

la. Ibid., Chap. 14. 

2. Influence of Burning Time, Mass Velocity, ayid 

Tube Wall Thickness on the Heat Failure of Rocket 
Tubes, Motor Test Section, OEMsr-418, Report 
IAC 6, California Institute of Technology, Jan. 20, 
1943. Div. 3-410-MI 

3. Machinery’s Handbook, p. 335. 

4. Rocket Torpedo Projects, James B. Edson, OEMsr- 
418, Report INC 2, California Institute of Tech- 
nology, Oct. 27, 1942. 

5. NDRC Armor and Ordnance Report No. A-166. 

6. Design of the High-Velocity Rocket (VICAR) 

(Final Report), John Beek, Jr., R. J. Thompson, 
G. D. Brewer, and R. R. Newton, OSRD 5820, 
OEMsr-273, ABL Report W-21, Service Projects 
W-210 and OD-201, George Washington University, 
1946. Div. 3-551.5-MI 

7. Small-Caliber High-Velocity Rocket (CURATE) 
(Final Report), R. J. Thompson, G. D. Brewer, 
and R. R. Newton, OSRD 5820, OEMsr-273, ABL 
Report W-21.1, Service Projects OD-201 and W-210, 
George Washington University, January 1946. 

Div. 3-551.4-MI 

8. Rocket Motor Tube Bending Machine for Decreas- 
ing CG Malalignment, L. A. Richards, OEMsr-418, 
Report IGC 4, California Institute of Technology, 
Mar. 1, 1943. 

9. Manual: Methods of Manufacture for the 4.5-in. 
BR, Lowell Martin, OEMsr-418, Report JSC 1, Cali- 
fornia Institute of Technology, Sept. 22, 1943. 

Div. 3-520-M10 


10. Study of Nozzle Side Forces by Means of Com- 

pressed Air Jet, Gabriel E. Kron and O. C. Wilson, 
OEMsr-418, Report ILC 1, California Institute of 
Technology, Dec. 15, 1942. Div. 3-440-M3 

11. Further Investigations Conducted with the Yaw 

Machine, Accuracy Committee, OEMsr-418, Report 
ILC 2, California Institute of Technology, Jan. 4, 
1943. Div. 3-614-MI 

12. Weekly Progress Report, OEMsr-418, Report PMC 
1.39, California Institute of Technology. 

Div. 3-110-M5 

13. Weekly Progress Report, OEMsr-418, Report PMC 
1.18, California Institute of Technology. 

Div. 3-110-M2 

14. Methods for One-Piece Nozzle Manufacture, De- 

velopmental Engineering Section, OEMsr-418, Re- 
port JSC 4, California Institute of Technology, 
Mar. 8, 1944. Div. 3-440-M4 

15. Supplement No. 1 : Methods for One-Piece Nozzle 
Manufacture, Developmental Engineering Section, 
OEMsr-418, Report JSC 4, Sup. 1, California In- 
stitute of Technology, May 16, 1944. Div. 3-440-M5 

16. Gas-Malalignment and Deflection-Malalignment 

Ratio for All Types of BR Fired from September 
20, 1942 to April 1, 1943, C. W. Snyder, OEMsr- 
418, Report IBC 23, California Institute of Tech- 
nology, April 1943. Div. 3-520-M8 

17. Weekly Progress Report, OEMsr-418, Report PMC 
2.27, California Institute of Technology. 

Div. 3-110-M9 

18. Weekly Progress Report, OEMsr-418, Report PMC 
2.4, California Institute of Technology. 

Div. 3-110-M5 

19. Weekly Progress Report, OEMsr-418, Report PMC 
2.8, California Institute of Technology. 

Div. 3-110-M5 

19a. Ibid., p. 12. 

20. Weekly Progress Report, OEMsr-418, Report PMC 
1.98, California Institute of Technology. 

Div. 3-110-M5 

21. Weekly Progress Report, OEMsr-418, Report PMC 
2.6, California Institute of Technology. 

Div. 3-110-M5 

22. A Study of Nozzle Erosion, Motor Test Group, 

OEMsr-418, Report IGC 7, California Institute of 
Technology, Mar. 8, 1944. Div. 3-441-MI 

23. Nozzle Erosion in the 3MR3 Rocket Motor Deter- 
mined from Static Firing Records, N. U. Mayall, 
OEMsr-418, Report IAC 13, California Institute of 
Technology, June 22, 1944. 


BIBLIOGRAPHY 


363 


24. Weekly Progress Report, OEMsr-418, Report PMC 
2.62, California Institute of Technology. 

Div. 3-110-M9 

25. Weekly Progress Report, OEMsr-418, Report PMC 
2.69, California Institute of Technology. 

Div. 3-110-M9 

26. A Study of Materials for Jet Motor Exhaust 
Nozzle, M. M. Mills, Report 18, Galcit Project 1. 

27. Nozzle Erosion in the 3MRS Rocket Motor Deter- 

mined from Static Firing Records, N. U. Mayall, 
OEMsr-418, Report IAC 14, California Institute of 
Technology, Aug. 7, 1944. Div. 3-441-M3 

28. Drag Characteristics of Various Aircraft Rocket 
Projectiles, Hsue-Shen Tsien and Leverett Davis, 
Jr., OEMsr-418, Report IPC 52, California Institute 
of Technology, Jan. 15, 1944. Div. 3-243.3-MI 

29. Closures and Seals for Rocket Motors, Bruce H. 
Sage, OEMsr-418, Report IAC 19, California In- 
stitute of Technology, Nov. 19, 1945. 

Div. 3-412-M4 

Y 

Chapter 24 

1. The Exterior Ballistics of Rockets, Leverett Davis, 
Jr., J. W. Follin, Jr., and Leon Blitzer, monograph 
(unclassified) to be published by the McGraw-Hill 
Book Company, Inc. 

la. Ibid., Sec. 2.24. 

lb. Ibid., Chap. 5. 

lc. Ibid., Chap. 3. 

2. Curves for External Ballistic Calculations on Low- 
Velocity Rockets Fired at High Angles, Leverett 
Davis, Jr., OEMsr-418, Report JPC 2.2, California 
Institute of Technology, Aug. 20, 1943. 

Div. 3-221-M3 

3. NDRC Report A-231. 

4. Deceleration Coefficient of the 2-in. AA at High 

Velocities, Leverett Davis, Jr., OEMsr-418, Report 
IPC 23, California Institute of Technology, Apr. 
22, 1943. Div. 3-246-MI 

5. Drag Characteristics of Various Aircraft Rocket 

Projectiles, Hsue-Shen Tsien and Leverett Davis, 
Jr., OEMsr-418, Report IPC 52, California Institute 
of Technology, Jan. 15, 1944. Div. 3-243.3-MI 

6. The Trajectories of Target Rockets, A. J. Dempster, 

OSRD 345, NDRC Report A-28, Service Project 
OD-26, Jan. 27, 1942. Div. 3-625-M2 

7. Lehrbuch der Ballistik, G. Cranz, Vol. I, J. 
Springer, 1935 (Edwards Bros., 1943). 

8. Rocket Fundamentals, OSRD 3992, OEMsr-273, 

ABL Special Report 4, George Washington Uni- 
versity, 1944. Div. 3-210-M3 

8a. Ibid., Chap. 4. 


9. Initial Conditions for the Calculation of the CWG 
Trajectories, Leverett Davis, Jr., OEMsr-418, Re- 
port MTC 2, California Institute of Technology, 

Mar. 21, 1942. Div. 3-572-MI 

/ 

10. Tip-Off about a Fixed Point , Wallace Hayes, 

OEMsr-418, Report OPC 3, California Institute of 
Technology, Nov. 17, 1944. Div. 3-490-M2 

11. Effect of Wind on the Mean Deflection of Rockets, 

Leon Blitzer, OSRD 1361, NDRC A-169, OEMsr- 
418, Report JPC 7, California Institute of Tech- 
nology, Feb. 1, 1943. Div. 3-243.1-MI 

12. Ammunition Manual for the U.5-in. BR (1100-yd), 

OEMsr-418, Report JBC 19, California Institute of 
Technology, July 26, 1943. Div. 3-852-MI 

13. Exterior Ballistic Tables, prepared by the Ordnance 
Department, U. S. Army, 1924, Vol. 1, Introduc- 
tion. 

14. Firing of Rockets from Aircraft: Launchers, 
Sights, Flight Tests (Final Report), R. V. Adams, 
C. D. Anderson, and others, OSRD 2549, OEMsr- 
418, California Institute of Technology, 1946. 

Div. 3-530-M4 

15. The Exterior Ballistics of Fin-Stabilized Aircraft 
Rockets, Leon Blitzer and Leverett Davis, Jr., 
OSRD 2528, OEMsr-418, Report JPC 24, Cali- 
fornia Institute of Technology, Aug. 20, 1945. 

Div. 3-561-M2 

16. Principles of Rocket Firing from Aircraft, Illus- 

trated, OSRD 2428, OEMsr-418, Report JNC 30, 
Service Project NO-170, California Institute of 
Technology, Apr. 2, 1945. Div. 3-530-M3 

17. A Theory for the Difference Between the True 

Angle of Attack and the Effective Angle of At- 
tack, Leverett Davis, Jr., OSRD 2527, OEMsr-418, 
Report JPC 30, California Institute of Technol- 
ogy, Sept. 28, 1945. Div. 3-245.2-MI 

18. The Exterior Ballistics of Fin-Stabilized Aircraft 
Rockets, Leon Blitzer and Leverett Davis, Jr., 
OSRD 2528, OEMsr-418, Report JPC 24, Cali- 
fornia Institute of Technology, Aug. 20, 1945. 

Div. 3-561-M2 

19. The Ballistics of Firing an ASB Backwards from 
an Airplane, Leverett Davis, Jr., OEMsr-418, Re- 
port ITC 4, California Institute of Technology. 

Div. 3-731. 1-M12 

20. Some Operational and Logistical Problems in the 

Use of Rockets, William A. Fowler, OSRD 2366, 
OEMsr-418, Report JNC 16, California Institute of 
Technology, Feb. 1, 1945. Div. 3-220-M4 

21. The Effect of Fin Size, Burning Time, and Pro- 
jector Length on the Accuracy of Rockets, Ira S. 
Bowen, Leverett Davis, Jr., and Leon Blitzer, 


364 


BIBLIOGRAPHY 


OSRD 1330, NDRC A-164, OEMsr-418, Report JPC 
3, California Institute of Technology, Jan. 4, 1943. 

Div. 3-249-M2 

22. Approximate Formulae for Calculation of Deflec- 
tion of Rockets, J. G. Waugh, OEMsr-418, Report 
IPC 45, California Institute of Technology. 

Div. 3-243. 1-M6 

23. Gas-Malalignment and Deflection-Malalignment 

Ratio for All Types of BR Fired from September 
20, 19^2 to April 1, 1943, C. W. Snyder, OEMsr- 
418, Report IBC 23, California Institute of Tech- 
nology, April 1943. Div. 3-520-M8 

24. Mechanical Destruction Tests of Metal Parts of 
Subcaliber Mk 1 ASPC Tails, Bruce H. Sage, 
OEMsr-418, Report IBC 24, California Institute of 
Technology, Apr. 20, 1943. 

25. I: Analysis of the Causes of Dispersion of the 

4.5-in. Barrage Rocket: II: Dispersion Data on 
CIT Rockets, OSRD 1632, NDRC A-73M, OEMsr- 
418, Report JPC 8, California Institute of Tech- 
nology, May 17, 1943. Div. 3-245-M2 

26. Supplement: I, Analysis of the Causes of Disper- 
sion of the 4.5-in. Barrage Rocket; II, Dispersion 
Data on CIT Rockets, OEMsr-418, Report JPC 8.2, 
California Institute of Technology, June 11, 1943. 

Div. 3-245-M2 

27. Experimental Attempts to Improve the Accuracy 

of Rockets, 0. C. Wilson and Gabriel E. Kron, 
OEMsr-418, Report JPC 9, California Institute of 
Technology, Nov. 26, 1943. Div. 3-249-M4 

28. Weekly Progress Report, OEMsr-418, Report PMC 

2.38, Part 2, California Institute of Technology, 
June 25, 1944, p. 7. Div. 3-110-M9 

29. Spiral Launching of 4.5-in. Rockets (Final Report) , 

R. R. Newton, OSRD 5813, OEMsr-273, ABL Re- 
port W-18.2, George Washington University, De- 
cember 1945. Div. 3-491. 237-M9 

30. The Relationship between Dispersion in Firing 

from a Plane and from the Ground, Ira S. Bowen, 
OEMsr-418, Report TPC 2, California Institute of 
Technology, Dec. 25, 1942. Div. 3-245-MI 

31. Dispersion due to Malalignment of Fin-Stabilized 
Rockets in Forward Firing from Aircraft, L. Ivan 
Epstein, OSRD 2190, OEMsr-418, Report JPC 23, 
Service Projects OD-16, NO-170, and others, Cali- 
fornia Institute of Technology, Aug. 10, 1944. 

Div. 3-561. 1-M3 

32. Ammunition Dispersion of Long -Burning Unro- 
tated Rockets in Forward Firing from Airplanes, 
Leon Blitzer, OEMsr-418, Report IPC 39.2, Cali- 
fornia Institute of Technology, Apr. 10, 1944. 

Div. 3-245-M4 


33. Sources of Error and Dispersion in Forward Firing 
of Nonrotating Aircraft Rockets, Leon Blitzer and 
Leverett Davis, Jr., OEMsr-418, Report JPC 19, 
Service Projects OD-162, NO-164, and others, 
California Institute of Technology, Apr. 25, 1944. 

Div. 3-532.3-M5 

34. Principles of Rocket Design, Fowler and Lauritsen, 
Chap. 8, “The Underwater Behavior of Rockets.” 

35. Radius of Curvature of the Underwater Trajec- 

tory of a Rocket, Leverett Davis, Jr., OEMsr-418, 
Report IPC 54, California Institute of Technology, 
Mar. 17, 1944. Div. 3-711-MI 

36. Estimation of Drag and Nose-Lift Coefficients of 
Some Rockets Necessary to Give a Prescribed 
Underwater Behavior, Leverett Davis, Jr., and 
L. Ivan Epstein, OEMsr-418, Report IPC 57, Cali- 
fornia Institute of Technology, Apr. 10, 1944. 

Div. 3-712-M2 

37. Weekly Progress Report, OSRD 2477, OEMsr-418, 

Report PMC 2.84, California Institute of Tech- 
nology, June 24, 1945. Div. 3-110-M9 

38. Weekly Progress Report, OSRD 2486, OEMsr-418, 

Report PMC 2.85, California Institute of Tech- 
nology, July 8, 1945. Div. 3-110-M9 

39. Weekly Progress Report, OSRD 2505, OEMsr-418, 

Report PMC 2.87, California Institute of Tech- 
nology, Aug. 5, 1945. Div. 3-110-M9 

Chapter 25 

1. The Exterior Ballistics of Rockets, Leverett Davis, 
Jr., James W. Follin, Jr., and Leon Blitzer, mono- 
graph (unclassified) to be published by the Mc- 
Graw-Hill Book Company, Inc. 

la. Ibid., Chap. 9. 

2. The Effect of Aerodynamic Moments on the Motion 

of Spin-Stabilized Rockets during Burning, James 
W. Follin, Jr., OSRD 2531, OEMsr-418, Report 
JPC 27, California Institute of Technology, Sept. 
21, 1945. Div. 3-562.1-M2 

3. Theoretical Curves Showing Yaw, Orientation, and 

Deflection of 3.5-in. SSR, 5.0-in. HCSR, and 5.0-in. 
HVSR During Burning, L. Ivan Epstein, OEMsr- 
418, Report JPC 20 Sup., California Institute of 
Technology, Dec. 30, 1944. Div. 3-243.4-MI 

4. Calculation of Mallaunching of Spin-Stabilized 
Rockets, Leverett Davis, Jr., and J. G. Waugh, 
OSRD 2235, OEMsr-418, ‘ Report JPC 22, Service 
Projects NO-33 and NO-213, California Institute 
of Technology, Sept. 20, 1944. Div. 3-562. 1-MI 

5. Range Tables for Spin-Stabilized Rockets, James 
W. Follin, Jr., and P. W. Stoner, OSRD 2536, 
OEMsr-418, Report JPC 31, California Institute of 
Technology, Nov. 15, 1946, Table 2. 

Div. 3-562-M3 


BIBLIOGRAPHY 


365 


6. The Effect of Wind on Ground-Fired Spin-Sta- 
bilized Rockets During Burning, James W. Follin, 
Jr., OEMsr-418, Report IPC 76, California Institute 
of Technology, Apr. 23, 1945. Div. 3-243. 1-M5 

7. Notes on the External Ballistics of Rotating 
Rockets, Leverett Davis, Jr., OEMsr-418, Report 
JPC 18, Service Project NO-215, California In- 
stitute of Technology, Apr. 6, 1944. Div. 3-220-M2 
7a. Ibid., Figure 3. 

8. Local memorandum CIT/OBC 41.4, Apr. 30, 1945. 

9. Nation Bureau of Standards Report RHH: SBB: 
VI, 4/44. 

10. Field Testing of Rockets: Range Operations and 
Metric Photography (Final Report), W. N. Arn- 


quist, R. R. Cox, and others, OSRD 2547, OEMsr- 
418, California Institute of Technology, 1946. 

Div. 3-610-M7 

11. The Influence of the Magnus Moment on the Sta- 

bility of Rotating Projectiles, Leverett Davis, Jr. 
and James W. Follin, Jr., OSRD 2529, OEMsr-418, 
Report JPC 29, California Institute of Technology, 
Sept. 1, 1945. Div. 3-243-M3 

12. Principles of Rocket Design, Fowler and Lauritsen. 

13. Range Tables for Spin-Stabilized Rockets, James 
W. Follin, Jr., and P. W. Stoner, OSRD 2536, 
OEMsr-418, Report JPC 31, California Institute of 
Technology, Nov. 15, 1945. 



OSRD APPOINTEES 


Division 3 

DIVISION A b — 1940-1942 

Chief c 

Richard C. Tolman 

Deputy Chief 

Charles C. Lauritsen 

Hannah Markstein 
Duane Roller 
Edith Townes 

Member 

Ernest C. Watson 


Technical Aides 

Eliot B. Bradford 
John Elder 
Paul C. Fine 


SECTION H— 1940-1942 

Chief c 

Clarence N. Hickman 

Vice Chairman 

Ralph E. Gibson 

Technical Aides 

Alexander Kossiakoff 

Members 

Edwin U. Condon 
William D. Coolidge 
Arthur J. Dempster 

SECTION C— 1942 

Chief c 

John T. Tate 

Technical Aide 

Lorenz G. Straub 

DIVISION 3—1943 

Chief 

John T. Tate 
Charles C. Lauritsen d 

Technical Aides 

Alexander Kossiakoff 
Lorenz G. Straub 
Otto A. Wantuch 


Alvin G. Anderson 
Eliot B. Bradford 
Philip T. Kirwan 


Philip T. Kirwan 

George B. Kistiakowsky 
Ernest C. Watson 
Warren Weaver 


OSRD APPOINTEES (Continued) 


Members 

Alexander Ellett 
Ralph E. Gibson 
Clarence N. Hickman 
Frederick L. Hovde 

Ernest C. Watson 


E. P. Hubble 
George B. Kis'tiakowsky 
William N. Lacey 
Charles C. Lauritsen 


DIVISION 3-1943-1946 

Chief 

Frederick L. Hovde 
Eliot B. Bradford* 

Technical Aides 

Eliot B. Bradford Grace L. Hart 

Frederick W. Cummings Roger S. Warner 

Philip T. Kirwan 


Members 

Alexander Ellett 
Ralph E. Gibson 
Clarence N. Hickman 

Louis P. Hammett 


George B. Kistiakowsky 
Charles C. Lauritsen 
Ernest C. Watson 


SECTON H— 1943-1946 

Chief 

Clarence N. Hickman 

Vice Chairman 

Ralph E. Gibson 

Technical Aides 

Philip T. Kirwan Julius A. Folse 

Glenn H. HoppiN e 

SECTION L— 1943-1946 
Frederick L. Hovde* 

Technical Aide 

Bayes M. Norton 


a Many of the appointments were in effect for periods 
shorter than those shown for organizational units. The 
listing is primarily in order of appointment dates. Si- 
multaneous appointees are arranged alphabetically. 

Certain short-lived organizations are not shown 
above. 


b Division A (Armor and Ordnance) included Sec- 
tions C and H on rocket development, and Sections A, 
B, E, and T in other fields. 

c Called Chairman through 1942. 
d Acting Chief. 

e Changed to Special Assistant in early 1946. 


367 


CONTRACT NUMBERS, CONTRACTORS, AND SUBJECT 
OF CONTRACTS 

The following list includes all contracts under which the rocket 
research and development programs of Division A and Division 3 
were carried out. In addition, there were two purchase con- 
tracts with the Hercules Powder Company for early supplies 
of rocket propellant, a transfer of funds to the Army Ordnance 
Department for the same purpose, and transfers of 1941, 1942, 
and 1943 funds to the Navy Bureau of Ordnance for support of 
the Jet Propulsion Research Laboratory at the Naval Powder 
Factory, Indian Head, Maryland. Contract OEMsr-418 (which 
included OEMsr-250) was the only contract under Sections C 
and L; all others (except OEMsr-673) were related to Section H 
programs. 

The scope of the work under each contract is indicated briefly 
below, and more completely by the report titles listed under each 
contract in Appendix Q. 


Contract Number 


Name and Address of Contractor 


Subject 


OEMsr-250 

California Institute of Technology 

Pasadena, California 

This contract included in and 
superseded by OEMsr-418. 

OEMsr-256 

Western Electric Company 

Bell Telephone Laboratory, Inc. 

New York, New York 

Instrumentation for measur- 
ing rocket performance. 
Development of rocket 
launchers, of firing mech- 
anisms, of components for 
Army M8 type 4.5-incn 
rockets, of propeller actu- 
ated ignition devices for 
bomb accelerators and of 
an electromagnetic fuze. 

OEMsr-273 

The George Washington University 
Washington, D. C. 

(with operations there, at Naval 

Powder Factory, Indian Head, 
Maryland, and at Allegany Ballis- 
tics Laboratory near Cumberland, 
Maryland) 

Central laboratory opera- 
tions. Ballistics research. 
Development and testing of 
instrumentation, improved 
propellants, flame throw- 
ers, improved mortars, re- 
coilless mortars, and many 
types of rockets and re- 
lated equipment. 

OEMsr-416 

Hercules Powder Company 

Wilmington, Delaware 
(work at Kenvil, New Jersey) 

Early improvements in rocket 
propellants of solvent-ex- 
truded ballistite types. 

OEMsr-418 

California Institute of Technology 

Pasadena, California 

(with operations there at Camp 

Haan, at Morris Dam, at Camp 
Pendleton, at Naval Ordnance Test 

Station, all in California) 

Central laboratory opera- 
tions. Ballistics research. 
Research and development 
on aircraft torpedoes and 
other underwater ordnance. 
Development of instru- 
mentation, of dry extru- 
sion, of ballistite propel- 
lant, of all rockets, most 
launchers and most of the 
rocket fuzes used by the 
U. S. Navy in World War 
II. Pilot production of these 
items. 


368 


CONTRACT NUMBERS, CONTRACTORS, AND SUBJECT OF CONTRACTS ( Continued ) 


Contract Number 

Name and Address of Contractor 

Subject 

OEMsr-671 

Budd Induction Heating, Inc. 

Detroit, Michigan 

Engineering designs and ex- 
perimental production of 
4 ".5 rockets. 

OEMsr-673 

Armour Research Foundation 

Chicago, Illinois 

Included no work on rockets. 
Transferred to Division 6 
in late 1943. 

OEMsr-702 

California Institute of Technology 

Pasadena, California 

Special studies of double 
base powders (work con- 
tinued under Division 8 
contract OSMsr-881). 

OEMsr-716 

University of Minnesota 

Minneapolis, Minnesota 

Studies of the burning of 
double-base propellants. 

OEMsr-733 

Duke University 

Durham, North Carolina 

Closed chamber studies of 
propellants (this contract 
was taken over from Divi- 
sion 1). 

OEMsr-762 

University of Wisconsin 

Madison, Wisconsin 

Burning of double-base pro- 
pellants. 

OEMsr-947 

Catalyst Research Corporation 

Baltimore, Maryland 

Development of gasless delay 
elements for ejection 
charges. 

OEMsr-968 

• 

Budd Wheel Company, Inc. 

Detroit, Michigan 

Engineering design and ex- 
perimental production of 
metal components for many 
rockets and mortars. 


369 



SERVICE PROJECT NUMBERS 


The projects listed below were transmitted to the Executive 
Secretary, National Defense Research Committee, NDRC, from 
the War or Navy Department through either the War Depart- 
ment Liaison Officer for NDRC or the Office of Research and 
Inventions (formerly the Coordinator of Research and Develop- 
ment), Navy Department. 


Service 

Project 

Number 


Subject 


AC-52 

AC-70 

AC-121 

CWS-10 

CWS-22 

CWS-30 

CWS-34 

NA-167 

NA-197 

NA-231 

NO-33 

NO-34.1 

NO-34.2 

NO-34.3 

NO-35.1 

NO-35.2 

NO-36.5 

NO-39.1 

NO-99 

NO-116 

NO-118 

NO-120 

NO-121 

NO-140 

NO-141 

NO-146 

NO-148 

NO-153 

NO-164 

NO-165 

NO-168 


Development of a specially shaped bomb (referred to as a water plunge bomb) de- 
signed to follow a horizontal path in water after being dropped at high speed 
from aircraft. 

Hydrobomb (torpedo for Army aircraft). 

Development of sights for firing aircraft rockets. 

Development of flame throwers (including their pressuring by propellant gases). 

Rocket projection of chemical munitions (and extension of range of 4 ".2 chemical 
mortar) . 

Development of 4".2 recoilless mortar and shell. 

Improvement of 4".2 chemical mortar. 

Study of nozzle design for jet motors. 

Development of jet-assisted take-off unit for carrier based aircraft. 

Assistance on the development of aircraft launching equipment. 

Internal and external ballistics of rockets; and double-base propellants for rockets. 
Rockets for aircraft armament. 

Low altitude antiaircraft rockets. 

High altitude antiaircraft rockets. 

Jet accelerators for armor-piercing bombs. 

Rockets for assisted airplane take-off (including rocket catapult). 

Rocket projection of antisubmarine depth bombs from ships. 

Rocket targets. 

Jet propulsion (solventless extrusion at Bruceton). 

Scatter bombs for attack of submarines by airplanes. 

Rocket weapons (for beach barrage in amphibious assault). 

Parachute rocket flare (for identification of warfare targets from aircraft). 
Retro-rocket bombs (for attack of submarines by MAD-equipped airplanes). 
Horn-type retro-bombing fuze. 

Hydrodynamic characteristics of projectile forms. 

Underwater trajectories of depth charges. 

Torpedo launching mechanism (design and construction of). 

Smoke float rocket and projector, development of. 

3 ".25 rocket and projector. 

Rocket projectors (for Marine Corps, development of, and establishment of test 
ranges) . 

Rocket deceleration of aircraft launched torpedoes. 


370 



SERVICE PROJECT NUMBERS ( Continued ) 


Service 

Project Subject 

Number 


NO-170 

NO-176 

NO-177 

NO-192 

NO-196 

NO-204 

NO-205 

NO-214 

NO-215 

NO-216 

NO-227 

NO-228 

NO-230 

NO-238 

NO-245 

NO-246 

NO-247 

NO-248 

NO-249 

NO-250 

NO-251 

NO-252 

NO-253 

NO-254 

NO-256 

NO-259 

NO-260 

NO-271 

NO-280 

NO-282 

NO-284 

NO-289 

NO-296 

NS-164 

NS-211 

NS-309 

OD-14 

OD-26 

OD-66 


OD-98 

OD-125 


Adaptation of Navy 3". 25 and 5".0 rockets to aircraft (and development of 5".0 high 
velocity aircraft rockets). 

Torpedoes for high speed aircraft (including water entry tests). 

Jet-propulsion of aircraft torpedoes. 

Shipboard rocket launcher for the 3". 25 rocket, development of. 

Anti-surface vessel ordnance. 

Contact fuzes, development of (for depth bombs). 

Rocket targets (production of). 

Ballistic range converter for ASD radar (for forward firing aircraft rockets). 

Spin stabilized rockets. 

Aircraft sight for forward firing rockets of CIT 3A type. 

Subcaliber training rocket for aircraft use. 

Design, construction, and operation of extrusion presses (at Naval Ordnance Test 
Station, Inyokern, Calif.). 

3 ".25 window rockets. 

Development of launchers for spin stabilized rockets. 

Development of high performance 4 ".5 aircraft rocket. 

Development of 5".0 aircraft rocket. 

Development of 2".36 high velocity rocket, H.E.A.T. 

Development of improved components for 4".5 rocket, M8 type. 

Development of spin-stabilized rocket using solvent-extruded propellants. 

Rocket projection of bombs. 

Development of 3 ".25 or 3 ".5 rocket. 

Development of 3 ".25 rocket, multiple grain, thin web. 

Development of 7 ".2 rocket motor. 

Development of 10" rocket motor. 

Forward firing large caliber aircraft rockets (“Tiny Tim”). 

Demolition rockets and launchers. 

Scoring of air to air rocket firing. 

Experimental production of 3". 25 rocket motors, Mk 5, for CWR-N rockets. 
Statistical assistance on the analysis of firing data for rocket propellant. 
Development of 2,000 lb forward firing, large caliber aircraft rocket. 

Development of aircraft rocket sights. 

Assistance to the Naval Ordnance Test Station, Inyokern, Calif. 

Development and fabrication of launching rockets for Bumblebee. 

Rocket propulsion to insure proper ejection of Mk 2 grenade from new airless emer- 
gency signal ejector of submarine. 

Countermeasure for antisubmarine contact fuzed charges. 

3".0 solid slow burning propellants (for generating gases to drive turbine). 

Special fuels for jet propulsion and squib igniter performance. 

Jet propulsion (development of many early rockets; superseded by OD-161 to 172). 

Device to determine direction and range of a forward artillery officer from immediate 
vicinity of a battery position. 

Rocket targets (with wings). 

Long range (75 miles) rocket projectile. 


371 



SERVICE PROJECT NUMBERS ( Continued ) 


Service 

Project Subject 

Number 


OD-137 

OD-155 

OD-161 

OD-162 

OD-163 

OD-164 

OD-165 

OD-166 

OD-167 

OD-168 

OD-169 

OD-170 

OD-171 

OD-172 

OD-179 

OD-183 

OD-184 

OD-185 

OD-186 

OD-187 

OD-196 

OD-199 

OD-201 


Demolition rockets and launchers. 

Factors which control afterburning in rockets. 

Development of high performance 4". 5 aircraft rocket. 

Development of 5" aircraft rocket. 

Development of 2". 36 high velocity rocket H.E.A.T. (and of electromagnetic fuze 
for it). 

Development of 3". 25 rocket, single grain, solventless powder type. 

Development of improved components for 4 ".5 rocket, Mk 8 type. 

Development of spin-stabilized rockets using solvent extruded propellant. 

Development of spin-stabilized rockets using solventless-extruded propellant. 

Rocket projection of bombs (by standard rocket motors). 

Development of 3 ".25 or 3 ".5 rocket. 

Development of 3". 25 rocket, multiple grain, thin web. 

Development of 7 ".2 rocket motor. 

Development of 10" rocket motor. 

Statistical assistance on the analysis of firing data for rocket propellant. 

Bourdon systems (for measurement of performance of rocket propellant grains). 

Development of powder charge assembly for recoilless mortar, 60 mm. 

Development of stationary rocket motor, 3 to 3.5 inch, for special H.E.A.T. projectile 
(includes 81 mm recoilless mortar). 

Minefield clearing devices of the jet-propelled type. 

Adaptation of Tiny Tim rocket motor. 

Multiple-cartridge, tube-launching system for JB-2. 

Rocket accessories for aircraft. 

Research on elements of rocket motors with high impulse ratio. 


372 


INDEX 


The subject indexes of all STR volumes are combined in a master index printed in a separate volume. 
For access to the index volume consult the Army or Navy Agency listed on the reverse of the half-title page. 


ABL (Allegany Ballistics Laboratory) 
internal burning rocket propellant 
grains, 247 

properties of rocket propellants, 
99-113 

rocket propellants, 93, 105 
thermodynamics of rocket propel- 
lants, 71-77 

Acceleration of rockets, 212-213 
Accelerometers for torpedo test mea- 
surements, 28-32 
Aerodynamic forces 
cross wind, 288, 298 
damping moment, 289, 298 
deceleration moment, 289, 298 
drag, 214-215, 270-271, 288, 298 
effect on fin-stabilized rockets, 268- 
269 

effect on spin-stabilized rockets, 
288-298 

Magnus force, 288-289, 298-301 
Air drag, effect on rocket trajectory 
ground-fired rockets, 270-271 
range, 214-215 

spin-stabilized rockets, 288, 298 
AIR rocket nose fuzes, 130-131, 175- 
176, 184 

Aircraft rockets, design, 126-127 
Aircraft rockets, launchers, 140-147 
damage to aircraft, 147 
design problems, 147 
displacement and drop launchers, 
144-146, 276 

effect on air trajectory, 275-276 

fixed-type, 144, 146 

for retro firing, 140-141, 167-168 

for spinners, 147 

Mark 4; 141, 175-176 

Mark 5; 142 

post launchers for forward firing, 
142-144, 275-276 

rail launchers for forward firing, 141, 
275-276 
tree-type, 143 
T-slot, 138, 141, 175 
Aircraft rockets, trajectory, 274-276 
angle of attack, 276 
comparison with bullets, 274-275 
control of underwater trajectory, 
127-128 
dispersion, 282 
effect of launchers, 275-276 
effect of wind, 272 


/ 

effect of yaw, 275 
range, 214-215 
spin-stabilized, 305 
velocity, 274-276 
Aircraft rockets, types 
3.25-in., 247, 249, 252 
3.5-in., 126, 128, 170-176, 217-218 
5.0-in. AR, 171, 175-176 
11.75-in., 186-195 
14-in., 248-249 
115-mm, 91 
antisubmarine, 5 
fin-stabilized, 122, 165-195, 282 
forward-firing, 141-144, 175, 218 
GP, 220 

high-velocity, 179-185 
Mule, 165 

spin-stabilized, 122, 147, 203-204 
subcaliber, 176-179 
Aircraft torpedoes, 13 
Alden Hydraulics Laboratory, 8 
Alkali nitrate for rocket propellants, 
107 

Allegany Ballistics Laboratory 
internal burning rocket propellant 
grains, 247 

properties of rocket propellants, 
99-113 

rocket propellants, 93, 105 
thermodynamics of rocket propel- 
lants, 71-77 

Ammonium perchlorate-asphalt rocket 
propellants, 106-107 
Ammonium picrate for rocket pro- 
pellants, 107 

Aniline for liquid rocket propellants, 67 
Antiaircraft training, target rockets 
see Target rockets for antiaircraft 
training 

Antisubmarine bombs, 3-8 
facilities for testing underwater per- 
formance, 5 
fast-sinking, 3 
hedgehog, 3-4, 148 
ordnance problem, 3 
retro bombs, 5, 140-141 
testing laboratory, 4 
Antisubmarine rockets, 148-151 
aircraft, 5 

designation and types, 149 
fuzes, 135, 167 
general shape, 150 
head shape, 127 


igniters, 150 
launchers, 149 
nozzle, 149 
propellant grain, 150 
related rockets, 151 
retro rockets, 140, 165-170 
specifications, 148-149 
subcaliber, 163 
tail, 150 
yaw, 218 

Antitank grenade 
factors affecting performance, 162 
motor design features, 162-163 
okra grain, 162 
specifications, 162 
AR 

see Aircraft rockets 

Asphalt-ammonium perchlorate rocket 
propellant, 106-107 

Asphalt-potassium perchlorate rocket 
propellant 

burning properties, 106 
manufacturing process, 106 
mechanical properties, 106 
thermodynamic properties, 106 
ASR 

see Antisubmarine rockets 

ATG (antitank grenade) 

factors affecting performance, 162 
motor design features, 162-163 
okra grain, 162 
specifications, 162 

Atomic bomb test, use of retro rocket 
motors, 168 

Automatic rocket launchers 
Mark 7; 154-155 
Mark 51; 204-207 

Ballistic studies, underwater, 8-12 
factors affecting model behavior, 
11-12 

scaled models, 8-10 

water entry of projectiles, 10 

Ballistics of fin-stabilized rockets 
see Fin-stabilized rockets, exterior 
ballistics 

Ballistics of rocket propellants, 39-51, 
96-98 

see also Fin-stabilized rockets, ex- 
terior ballistics; Spin-stabilized 
rockets, exterior ballistics 
burning characteristics, 39, 41-45, 
96-97 


■■■tt Bute 


374 


INDEX 


charge design, 47-49 
discharge coefficient, 96 
drag of gas stream, 98 
effects of acceleration, 45-46 
heat transfer to the motor walls, 98 
internal-burning grains, 50 
liquid fuels, 40, 50-51 
nonsteady-state rockets, 98 
practical limitations, 40-41 
pressure, 96 

principles of propulsion, 39 
radiation, 97 
recommendations , 49-50 
resonance effect, 98 
specifications, 44-45, 98 
temperature limits , 46-47 
throat-to-port ratio, 96-97 
Ballistics of spin-stabilized rockets 
see Spin-stabilized rockets, exterior 
ballistics 

Ballistite for rocket propellants, 118, 
170, 187 

Barlow’s formula for wall thickness of 
rocket motors, 244 
Barrage rockets, 151-156 
4.5-in., 274 
accuracy, 153-154 
designation and types, 151-152 
fast-burning, 156 
for detonating land mines, 156 
heads, 153 

launchers and service use, 154-155 

military requirements, 151 

motor, 152 

stability, 167, 220 

tails, 153 

yaw, 218 

Bell Telephone Laboratories, 65 
Black powder rocket igniters, 52, 240 
Boat rocket launcher, 204 
Bombs, antisubmarine 
fast-sinking, 3 
hedgehog, 3-4, 148 
retro, 5, 140-142, 165-170, 252 
BR 

see Barrage rockets 
Brass can rocket igniters, 241 
British depth bomb (hedgehog), 3-4, 
148 

Burning strand method of studying 
rocket propellants, 82-83, 101 

Cant angle of rockets, 216 
Cast double-base rocket propellants 
advantages, 105 
process, 104-105 
recommendations, 110 
Cast perchlorate rocket propellants, 
105-107 

advantages, 105-106 
asphalt-ammonium, 106-107 


asphalt-potassium, 106 
ethylcellulose-potassium, 107 
general description, 105 
manufacturing process, 105-106 
nominal compositions, 105 
recommendations, 111 
Cavitation, torpedo, 17 
Cellulose acetate, use in rocket pro- 
pellants, 50, 60 

Centralite for rocket propellants, 83, 
102 

Chemical spinner, 203 
Chemical warfare grenade 

factors affecting performance, 162 
motor design features, 162-163 
okra grain, 162 
specifications, 162 
Chemical warfare rockets, 156-158 
accuracy, 158 
designation and types* 158 
dispersion, 279-280 
fuze, 158 

launchers and service use, 158 
motor, 156, 168 
propellant grain, 158 
spinner rocket, 203 

Chromium trioxide for rocket propel- 
lants, 106 

Chuffing of rocket motors, 235 
Closed bomb for studying rocket pro- 
pellants, 81 

Closed-breech rocket launcher, 169-170 
Colloidal rocket propellants 

see Double-base rocket propellants 
Composite rocket propellants 
composition, 68 
molded, 107-109, 111 
recommendations, 111 
solvent-extruded, 108-109, 111 
Cordite for rocket propellants, 170 
Crate rocket launcher, 154 
Cross force, effect on rocket’s under- 
water trajectory, 282-283 
Cross wind, effect on spin-stabilized 
rockets, 288, 298 
Cruciform rocket propellant grains 
advantages, 238-239 
applications, 234 
ballistite charge, 187 
upper temperature limit, 170-171 
CWG (chemical warfare grenade) 
factors affecting performance, 162 
motor design features, 162-163 
okra grain, 162 
specifications, 162 
CWR 

see Chemical warfare rockets 
CWSR (chemical warfare spinner rock- 
et), 203 



Damage instruments for torpedo test 
measurements, 32 

Damping moment, effect on spin- 
stabilized rockets, 289, 298 
DDR rocket base fuzes, 133-134 
Deceleration coefficient of rockets, for- 
mula, 214-215 

Deceleration moment, effect on spin- 
stabilized rockets, 289, 298 
Demolition rocket, 151 
Depth and roll recorder, 36 
Depth bombs, 3-4, 148 
Diethylene glycol dinitrate for rocket 
propellants, 69, 102 
DINA (explosive plasticizer) for rocket 
propellants, 102 

Dinitrotoluene for rocket propellants, 
69, 102 

Diphenylamine for rocket propellants, 
61, 69 

Discharge coefficient of rocket pro- 
pellants, 72-74 
formulas, 72, 96, 226 
throat-to-port ratio, 96 
Displacement rocket launchers, 144- 
146 

Double-base rocket propellants, 102- 
105 

burning properties, 83, 85, 102-103 
cast type, 104-105, 110 
composition, 68-69, 103-104 
granulations, 103-105 
mechanical properties, 103 
nitroglycerin, 69, 102 
recommendations, 109 
solvent-extruded, 103, 110 
solventless, 104, 110 
T-2; 43-44, 79-80 

Double-base rocket propellants, dry- 
processed, 56-63 
extrusion of stock, 57-58 
inhibiting of grain, 60 
machining, 59 

manufacturing process, 56-57 
recommendations, 62-63 
stability, 61-62 
types, 56 

DR (demolition rocket), 151 
Drag, effect on rocket trajectory 
ground-fired rockets, 270-271 
range, 214-215 

spin-stabilized rockets, 288, 298 
Drag coefficient, torpedo, 16-17 
Drag ring for torpedoes, 16, 33-35 
Drift signal rockets, 169-170 
Drop rocket launchers, 144-146, 276 
Dry-processed double-base rocket pro- 
pellants 

see Double-base rocket propellants, 
dry-processed 
Duke University, 81 


INDEX 


375 


11.75-in. aircraft rocket, 186-195 
blowout disk, 190, 260 
charge support, 190-192 
design problems, 186 
effect of firing temperature, 246 
fuzes, 194 
grid, 190 
head, 186, 194 
igniters, 192-194 
launchers, 144-147, 195 
lug bands, 189-190 
motor, 187-188, 193, 247, 248-249 
nozzle plate, 188 
propellant, 187 
tails, 189 

types and designations, 194 
use, 186 
Ethyl cellulose 

inhibiting coatings for rocket pro- 
pellants, 50, 92-93 
lacquer for rocket motor walls, 249 
Ethyl centralite, stabilizer for rocket 
propellants, 44-45, 61, 69 
Ethylcellulose-potassium perchlorate 
rocket propellant, 107 

False crimp rocket igniters, 241-242 
Fast-burning barrage rocket, 156 
Fast-sinking bombs, 3 
Fin-stabilized rockets, characteristics, 
121-123 

accuracy, 121-122 

comparison with spin-stabilized, 121- 
123 

head, 126, 127 
internal-burning grains, 50 
payload, 122 

simplicity and cheapness, 122 
tail, 262 

underwater stability, 122 
velocity, 118 
versatility, 122 

Fin-stabilized rockets, dispersion 
causes, 276-277 
fired from airplanes, 282 
ground firing, 278-282 
malalignment, 276-277 
suggestions for improved accuracy, 
280-282 

theory, 218-219 
yaw, 218-219 

Fin-stabilized rockets, exterior ballis- 
tics, 217-219, 267-287 
aerodynamic forces, 268-269 
air flight, 214-215, 274-276, 282 
ballistic quantities, 211 
center of pressure, 217 
comparison with spin-stabilized, 216, 
289, 306 

jet force and torque, 268-269 


range of a ground-fired rocket, 270- 
272 

retro firing, 276 
rocket motion, 267-268 
stable equilibrium, 217 
underground trajectories, 285-287 
underwater trajectories, 282-285 
wind effect, 272-274 
Fin-stabilized rockets for aircraft, 165- 
195 

2.25-in., 176-179 

3.5-in., 170-176 

5.0-in. high-velocity, 179-185 
11.75-in., 186-195 

comparison with spin-stabilized, 122 
dispersion, 282 
retro rockets, 165-170 
Fin-stabilized rockets for surface war- 
fare, 148-164 

antisubmarine, 5, 148-151 
barrage rockets, 151-156 
chemical warfare grenade, 162-163 
chemical warfare rockets, 156-158 
rocket grenade, 164 
subcaliber rockets, 163 
target rockets, 158-161 
Firing systems for rockets 

see Rockets, propulsion mechanism 

5.0-in. AR (aircraft rocket), 171, 175- 
176 

5.0-in. fin-stabilized rocket 

see High-velocity fin-stabilized rocket 

5.0-in. spin-stabilized rockets, 195-207 
aircraft spinners, 203-204 
heads, 200 

high-capacity spinners, 203 
launchers, 204-207 
Mark 7; 154-155, 172-173, 201-202, 
204 

pyrotechnic spinners, 203 
range, 200 

smoke and chemical spinners, 203 
spinner designations, 200-201 
Fixed rocket launchers, 144, 146 
Formulas for rockets 
acceleration, 212-213 
burning rate, 85, 96-97, 227 
burnt velocity, 39, 270 
cant angle, 216 

deceleration coefficient, 214-215 
discharge coefficient, 72, 96, 226 
effective gas velocity, 39, 211-212, 
213-214 

equilibrium pressure, 78-79, 96, 97, 
226 

linear rate of burning, 78, 226 
momentum, 211-212, 216 
overturning moment, 219 
range in free flight, 214 
specific impulse, 71-72 
stability factor, 219 



thrust coefficient, 74, 213-214 
vacuum range, 270 
velocity, 71 

wall thickness of motors, 244-245 
yaw of spinners, 221 
Forward-firing aircraft rockets 
launchers, 141-144, 175 
yaw, 218 

4.5-in. barrage rocket, 274 

4.5-in. spinner rocket, 91 
14-in. aircraft rocket, 248-249 
Foxboro depth and roll recorder, 36 
Fuels for rockets 

see Rocket propellants 
Fuzes for rockets, 129-137 
AIR nose fuzes, 130-131, 175, 184 
DDR base fuzes, 133-135 
general requirements, 129 
M48; 198 

Mark 139; 135, 167 
Mark 148; 175 
Mark 149; 131, 175, 184 
methods of arming, 129-130 
NIR nose fuzes, 131 
PIR base fuzes, 131-133 

G 117B rocket powder 
burning rate, 110 
pressure exponent, 79-80 
Galcit 61-C rocket propellant, 106 
Gas velocity in rocket propellants 
calculation of gas properties, 74-75 
control of rate, 101 
effect of temperature, 223-224 
effect on burning rate, 41 
effective velocity, 39, 99, 211-213 
theory, 71-72 

Gasoline for rocket propellants, 67 
GASR (general purpose spin-stabilized 
rocket), 220 

General purpose aircraft spin-stabilized 
rocket, 220 

George Washington University 
internal burning rocket propellant 
grains, 247 

properties of rocket propellants, 
99-113 

thermodynamics of rocket propel- 
lants, 71-77 

Granulation in rocket propellants, 
91-94, 103-105 
Grenades 

chemical warfare, 162-163 
incendiary rocket, 164 
Ground rocket launcher, 195 
Ground-fired rockets, range, 270-274 
air drag, 270-271 
calculation, 271 
dispersion, 278 
effect of burning time, 270 
launcher tip-off effects, 271-272 



376 


INDEX 


vacuum range, 270 
wind effects, 273-274 
Guggenheim Aeronautical Laboratory, 
68 

Guns, comparison with rockets 
efficiency, 123-125 
propellant grains, 100 
velocity, 274 

H-4 rocket propellant 
advantages, 42-44 
pressure exponent, 79-80 
H-5 rocket propellant, 83 
HCSR (high-capacity spinner rocket) 
ballistic constants, 289-290 
description, 203 

relation between velocity and pay- 
load, 118-119 

Hedgehog (British forward-thrown pro- 
jectile), 3-4, 148 

Hercules Powder Company, rocket pro- 
pellant grains, 94 
High-capacity spinner rocket 
ballistic constants, 289-290 
description, 203 

relation between velocity and pay- 
load, 118-119 

High-velocity fin-stabilized rocket, 179- 
185 

blowout disk, 260 
composition of steel in motor, 247 
dispersion, 278-279 
launchers, 185 

low-temperature performance, 180 
Mark 18 propellant grain, 179, 197 
maximum tubular propellant grain, 
238 

temperature distribution in motor 
wall, 246 
velocity, 118 

High-velocity fin-stabilized rocket, de- 
sign, 180-185 
fins, 182 
fuzes, 184 
heads, 183-184 
igniters, 182-183 
nonwelded motors, 184-185 
nozzle, 181 

seals and closures, 183-184 
suspension lugs, 182 
tubing, 180 
White Whizzer, 185 
High-velocity spinner rocket 
grain, 201 
heads, 202 
igniter, 201-202 
motor tube, 201-202 
nozzle plate, 202 
Holy Moses 

see High-velocity fin-stabilized rocket 


HVAR 

see High-velocity fin-stabilized rocket 
HVSR 

see High-velocity spinner rocket 
Hydrazine for rocket propellants, 67 
Hydro pressure plugs for torpedoes, 
33-35 

Hydrogen peroxide for rocket propel- 
lants, 67 

Igniters for rocket propellants, 52-55, 
239-243 

black powder, 52, 240 
construction and performance, 53-55 
containers, 54-55, 182-183, 192-193, 
241-243 

desirable characteristics, 241 
electric squibs, 54-55, 240 
function, 239-240 
Mark 17; 201-202 
Mark 18; 197 
principles, 52-53 
requirements, 55 
short ignition delays, 239-240 
Impulse of rocket propellants 

see Specific impulse of rocket pro- 
pellants 

IRG (incendiary rocket grenade), 164 

JP rocket propellant 

influence of position upon burning 
rate, 41 
stabilizer, 61 

JPH rocket propellant, 61 
JPN rocket propellant 
burning rate, 110 
impact energy, 44 
internal-burning grains, 50 
performance, 40 
specific impulse, 40, 45 
stability, 44, 61 

Kinetics of rocket propellants, 78-88 
area of burning surface, 80 
burning rate of powders, 80-87 
effect of powder composition, 80, 86- 
87 

pressure, 78-80 
rate of gas production, 78 
recommendations , 112-113 
theory of burning, 78-79, 87, 112 

L 4.8 rocket propellant 
burning rate, 110 
pressure exponent, 79-80 
rate of burning-pressure curves, 83 
temperature coefficient, 85 
Land mines, detonation by rockets, 156 
Launchers for rockets, 138-147 
airborne launchers, 140-147, 167-168, 
175-176, 275-276 



closed-breech launcher, 169-170 
crate launcher, 154-155 
mallaunching, 221, 294-296, 302-303 
Mark 6; 179 

Mark 7 (automatic launcher), 154- 
155 

Mark 20; 149 
Mark 22; 149 
Mark 40; 199 

Mark 50 (boat launcher), 204 
Mark 51 (automatic launcher), 204- 
207 

M-rail, 161 

post launcher, 142-144, 195, 275-276 
rail launchers, 138, 141, 161, 275-276 
seaborne launchers, 138-141, 204 
steel launcher, 185 
T-32; 158 
T-40; 151 

tip-off effects, 271-272 
Launchers for torpedoes 

see Torpedoes, launching tests 
Lift, effect on spin-stabilized rockets, 
288, 298 

Liquid rocket propellants 
advantages, 40 
application, 50-51 
requirements, 50 
types, 67 

M-8 rocket, 91 
M-48 rocket fuze, 198 
MAD (magnetic airborne detector), 
use of retro rockets, 5, 165 
Magnus force, effect on spin-stabilized 
rockets, 288-289, 298-301 
Mark I rocket propellant grain, 232- 
233 

Mark I torpedo drag ring 
effect on water entry, 16 
reduction of localized pressure, 33-35 
Mark 4 (T-slot) rocket launcher, 141, 
175-176 

Mark 5 rocket head, 183-184 
Mark 5 rocket launcher, 142 
Mark 6 rocket head, 167 
Mark 6 rocket launcher, 179 
Mark 6 rocket motor, 172 
Mark 7 high-velocity spinner rocket 
grain, 201 
heads, 202 
igniter, 201-202 
motor tube, 201-202 
nozzle plate, 202 
Mark 7 rocket launcher, 154-155 
Mark 7 rocket motor, 172-173 
Mark 8 rocket head, 203 
Mark 13 rocket propellant grain 
compressive stress on grain, 236 
effects of acceleration, 45-46 


INDEX 


377 


extrusion process, 58 
Mark 13 torpedo, 13-15 
design modifications, 15 
dive resistance, 18 
limitations, 13 
shroud ring tail, 14-15 
Mark 16 rocket propellant grain, 177, 
238 

Mark 17 rocket igniter, 201-202 
Mark 18 rocket igniter, 197 
Mark 18 rocket propellant grain, 179 
Mark 20 rocket launcher, 149 
Mark 21 rocket propellant grain, 201 
Mark 22 rocket launcher, 149 
Mark 23 rocket propellant grain, 
196-197 

Mark 25 torpedo, 15 
Mark 40 rocket launcher, 199 
Mark 50 ship rocket launcher, 204 
Mark 51 automatic rocket launcher, 
204-207 

Mark 139 rocket fuze, 135, 167 
Mark 148 rocket fuze, 175 
Mark 149 rocket fuze, 131, 175, 184 
Methyl alcohol for rocket propellants, 
67 

Methyl centralite for rocket propel- 
lants, 69 

Military requirements for rockets, 
117-125 
accuracy, 121 
barrage rockets, 151 
choice of fin or spin stabilization, 
121-123 

efficiency of rocket artillery, 123-125 
general characteristics and uses, 

117-118 

limitations, 125 
propulsion efficiency, 123-125 
range, 118 

underwater trajectory, 284-285 
velocity and payload, 118-121 
Mine clearance, use of retro rockets, 
168 

MJA rocket propellant, 85 
Molded composite rocket propellants, 
107-109 

general description, 107 
granulations, 108 
nominal compositions, 107 
properties, 108 
recommendations, 111 
Molded double-base rocket propellants 
process, 105 
recommendations, 110 
Molybdenum rocket nozzles, 257-258 
Momentum of rockets, formulas, 211- 
212, 216 

M orrisDamLaboratory , establ ishment , 4 
Mousetrap 

see Antisubmarine rockets 


M-rail rocket launcher, 161 
Mule (aircraft rocket), 165 

NIR rocket noze fuze, 131 
Nitrocellulose for rocket propellants 
effect on mechanical strength and 
elastic properties, 103 
instability, 44, 61 
preparation, 56-57 

Nitroglycerine, use in rocket propel- 
lants, 69, 102 

Nitro-methane for rocket propellants, 
67 

Nozzle design of rockets, 250-261 
accuracy, 250 
blowout disks, 260-261 
brazed-in formed nozzles, 178 
characteristics, 250, 258 
discharge coefficient, 226 
erosion, 256-259 
flash suppression, 255-256 
fuzes, 130-131, 175, 184 
materials, 257-259 
multiple nozzle, 173, 188, 253-255 
single nozzles, 251-253 
stellite nozzles, 258 
tolerances, 254-255 
types, 250-251 

Nutation in spin-stabilized rockets, 288 

Okra rocket propellant grains, 162, 234 
115-mm aircraft rocket, 91 
Orientation angle for rockets, 267 
Orientation curves for spin-stabilized 
rockets, 289-291 

Orientation recorders for torpedo test 
measurements, 35-36 
Overturning moment of rocket, 219, 
289-291 

Perchlorate rocket propellants, cast 
see Cast perchlorate rocket propel- 
lants 

Permafil (resin) for rocket propellants, 
107 

Photography of underwater torpedoes, 
27-28 

Phthalate esters for rocket propellants, 
102 

Pickle barrel (torpedo drag ring) 
effect on water entry, 16 
reduction of localized pressure, 33 
PIR rocket base fuzes, 131-133 
gas seals, 133 
method of arming, 131 
Pitch of torpedo , definition , 35 
Plastic case rocket igniters 
design, 182-183 
evaluation, 241 
Plastic rocket propellants 
manufacturing process, 109 



recommendations, 111 
Plasticizers for rocket propellants 
centralite, 83 

diethylene glycol dinitrate, 69, 102 
DI^TA, 102 
nitroglycerin, 69, 102 
triacetin, 83 

Post rocket launchers, 142-144, 195 
effect on trajectory, 275-276 
Mark 5; 142 
tree-type, 143 

Potassium nitrate for rocket propel- 
lants, 69 

Potassium perchlorate-ethylcellulose 
rocket propellant, 107 
Potassium salts for rocket propellants, 
69, 86-87, 102 

PySR (pyrotechnic spinner rockets), 
203 

Rail rocket launchers 
effect on trajectory, 275-276 
Mark 4; 141, 175-176 
M-rail, 161 
operation, 138 
T-32; 158 
Range of rockets 
deceleration coefficient, 214-215 
effect of burning time of propellant, 
270 

ground-fired, 270-272 
in air, 214-215 
in vacuum, 212-214, 270 
military requirements, 118 
spin-stabilized, 304-305 
Recommendations for future research 
ballistics of rocket propellants, 49-50 
cast double-base propellants, 110 
cast perchlorate propellants, 111 
chemistry of rocket propellants, 112, 
113 

dry-processed double-base rocket 
propellants, 62-63 

kinetics of rocket propellants, 112- 
113 

molded composite propellants, 111 
physical properties of rocket pro- 
pellants, 112 
plastic propellants, 111 
pressure molding of double-base 
powder, 110 

solid rocket propellants, 112 
solvent-extruded composite propel- 
lants, 111 

solvent-extruded double-base pow- 
ders, 110 

solventless double-base powders, 110 
Resonance effect in rocket propellants, 
98 

Retro rockets, 165-170 
design features, 167 



378 


INDEX 


designation and types, 165-166 
drift signal rockets, 169-170 
effectiveness 168 
launchers, 140-142, 167-168 
nozzle, 252 
related rockets, 168 
use in atomic bomb test, 168 
use in mine clearance, 168 
use with magnetic airborne detector, 
5, 165 

Ring tails, rocket, 261-262 
Ring tails, torpedo, 14-15 
Rocket fuzes 

see Fuzes for rockets 
Rocket grenade, 164 
Rocket heads, 126-128 
alignment, 126 
double-ogive, 127 
ground penetration, 285-287 
joint strength, 126 
leakage and heating, 126 
Mark 5; 183-184 
Mark 6; 167 
Mark 8; 203 
special shapes, 127-128 
zinc heads, 179 
Rocket motors 
3A9; 171 

Mark 6 (3A12), 172 
Mark 7 (3A16), 172-173 
White Whizzer, 185 
Rocket motors, design, 244-250, 262- 
266 

5.0-in. motor, 184-185 
chuffing, 234-235 

failures at high temperatures, 235- 
237 

for internal-burning grains, 247 
grain support, 263-264 
heating problems, 244-246 
insulation, 246-247 
internal pressure, 225-228 
performance calculations, 228-231 
reaction of wall with propellant, 249 
research and facilities required, 70 
seals, 194, 264-266 
straightness of tube, 249 
suspension lugs, 262-263 
threads, 247-249 
tube dimensions, 244 
tubing material, 244 
use of ethyl cellulose lacquer, 249 
wall thickness, 244-246 
weight, effect on velocity, 120 
weldability, 247 
Rocket orientation angle, 267 
Rocket performance, theory, 211-222 
fin stabilization, 217-219 
mechanism of propulsion, 211-214 
range, 214-215 

spin stabilization, 215-217, 219-222 


Rocket propellants, burning character- 
istics, 41-44, 80-87 
average rate, 42-44, 99 
burnt velocity, 39, 270 
composition and thermal properties, 
43 

effect on range, 270 
effect on total impulse, 45 
formulas, 85, 96-97, 227 
gas velocity, 41 
linear rate of burning, 78, 226 
position in grain, 41 
pressure, 42, 83-85, 100 
radiation, 86, 87, 97 
temperature, 42-44, 85 
theory, 78-79, 87-88, 112 
throat-to-port ratio, 96-97 
time of burning, 212-213 
Rocket propellants, burning rate meas- 
urements 

burning strand method, 82-83, 101 
closed bomb method, 81 
vented vessel method, 81-82 
Rocket propellants, characteristics, 96- 

102 

ballistic characteristics, 96-98 
function, 211 
gas temperature, 101 
gas velocity, 99, 101, 211-212, 

223-224 

mechanical properties, 102 
physical properties, 94, 112 
recommendations, 109-113 
resonance effect, 98 
sensitivity, 102 

specific impulse, 71-72, 99-101, 211- 
212 

stability, 101 

temperature, 223-225, 234-237, 244- 
246, 259-261 

thrust coefficient, 74, 211-214 
web thickness, 100 

Rocket propellants, design, 89-92, 
223-225, 239-243 
ballistic requirements, 47-49 
ballistite, 118, 170, 187 
catalyst, 106 
coolants, 69, 102 
desiccant bags, 243 
gasoline, 67 
grids, 243 

igniters, 52-55, 182, 192-193, 239-242 
insulation, 101 

low-temperature performance, 234- 
235 

maximum weight, 49 
oxidizers and fuels, 67 
Permafil, 107 
plasticizers, 69, 83, 102 
potassium salts, 69, 86-87, 102 
pressure-time curves, 225 


regressive type, 47 
silica gel, 243 
solventless process, 69 
specifications, 44-45, 98, 223 
stabilizers, 44-45, 61, 69, 102 
web thickness, 100 

Rocket propellants, grain characteris- 
tics, 236-239 
casting, 104-105 

comparison with gun propellant 
grains, 100 

effect on loading density, 101 
grain inhibitors, 50, 60 
internal-burning grain, 47-48 
length of grain, 48 
pressure molding, 105 
relation between shape and weight, 
238 

solvent extrusion, 94, 103 
solventless extrusion, 104 
stability requirements, 231-232 
stresses on grains, 236 
support, 263-264 
use of carbon dioxide, 94 
use of ethyl cellulose, 50, 92-93 
Rocket propellants, grain types, 91-94, 
231-234 

cruciform, 170-171, 234 
end-burning, 234 
inhibited, 92-94 
internal-burning, 50, 234, 247 
Mark 1; 232-233 
Mark 13; 45-46, 58, 171, 236 
Mark 16; 177, 238 
Mark 18; 179 
Mark 21; 201 
Mark 23; 197 
maximum weight, 237-239 
multiweb, 234 
okra, 162, 234 
single grain, 93 
tubular, 227, 231-233 
Rocket propellants, igniters 

see Igniters for rocket propellants 
Rocket propellants, internal pressure, 
78-80 

effect of temperature, 225, 228 
effect of throat-to-port ratio, 96 
effect on arming rocket fuze, 130 
effect on burning rate, 42, 83-85, 100 
effect on thrust coefficient, 213-214 
equilibrium pressure , 78-79 , 96-97 , 226 
resonance effect, 98 
Rocket propellants, theory 

see Ballistics of rocket propellants; 
Kinetics of rocket propellants; 
Thermodynamics of rocket pro- 
pellants 

Rocket propellants, types 
cast perchlorate, 105-107, 111 
composite, 68, 107-109, 111 



INDEX 


379 


double-base powders, 56-63, 69, 102- 
104, 110 

ethylcellulose-potassium perchlorate , 
107 

external-burning grains, 47 

G 117B powder, 79-80, 110 

Galcit 61-C, 106 

H-4; 42-44, 79-80 

H-5; 83 

JP, 41, 61 

JPH, 61-62 

JPN, 40, 44-45, 50, 61-62, 110 
L4.8; 80, 83, 85, 110 
liquid, 40, 50-51, 67 
MJA, 85 

nitrocellulose, 44-45, 56-57, 61, 103 
plastic, 109, 111 
Rocket tails, design 
fin tails, 262 
ring tails, 261-262 
Rockets, general types 

see also Fin-stabilized rockets; Spin- 
stabilized rockets 

barrage, 151-156, 167, 218, 220, 274 
chemical warfare grenade, 162-163 
chemical warfare rocket, 156-158, 
168, 203, 279-280 
demolition, 151 
drift signal, 169-170 
for underwater targets, 5 
forward-firing, 141-144, 175, 218 
ground-fired, 270-274, 278 
high-capacity spinner, 119, 203, 

289-290 

high-velocity fin-stabilized, 179-185 
high-velocity spinner, 201-202 
nonsteady-state, 98 
retro, 5, 140-141, 165-170, 252 
rocket grenade, 164 
ship-to-shore, 170 
smoke float, 168 
smoke spinner, 203 
subcaliber, 163, 176-179, 252 
target rockets, 158-161 
window rockets, 136-137, 168 
Rockets, launchers 

see Launchers for rockets 
Rockets, military requirements 

see Military requirements for rockets 
Rockets, nozzle design 

see Nozzle design of rockets 
Rockets, propulsion mechanism, 211- 
214 

burning time and acceleration, 212- 
213 

comparison with guns, 123-125 
components, 139-140 
effect of propellant temperature, 214 
efficiency, 123-125 

momentum - impulse - thrust , rela- 
tions, 211-212 


principle of operation, 67 
relation of pressure to thrust, 213- 
214 

solid fuel propulsion system, 48 
theory, 39 
Rockets, range 

see Range of rockets 
Rockets, specific models 

2.25- in. fin-stabilized, 53, 176-179, 
252 

3.25- in., 247, 249, 252 

3.5-in. fin-stabilized, 126, 128, 170- 
176, 217-218 

3.5- in. spin-stabilized, 196-199 
3R1; 196 

4.5- in. barrage rocket, 274 

4.5-in. spinner, 91 

5.0-in. high velocity, 179-185 
5.0-in. spin-stabilized, 154-155, 172- 
173, 195-207 

11.75-in. aircraft rocket, 186-195, 
248-249 

14-in. aircraft rocket, 248-249 
115-mm; 91 
M-8; 91 

Mark 7; 201-202 
T-59 (superbazooka), 91 
Vicar, 93 
Rockets, stability 

see Stability of rockets; Yaw of 
rockets 

Rockets, trajectory 

see Trajectory of rockets 
Rockets, velocity 
see Velocity of rockets 
Rockets for antiaircraft training 

see Target rockets for antiaircraft 
training 

Roll and depth recorder, 36 
SCAR 

see Subcaliber aircraft rocket 
Ship rocket launchers, 138-140 
blast, 139 

firing systems, 139-140 
Mark 50; 204 
types, 138-139 

Ship-to-shore rocket, specifications, 170 
Shroud ring tails, torpedo, 14-15 
Silica gel for rocket propellants, 243 
Slot rocket launchers, 138 
Smoke float rocket, 168 
SmSR (smoke spinner rocket) , 203 
Solid rocket propellants 
see Rocket propellants 
Specific impulse of rocket propellants 
burning at low pressures, 100 
definition, 39, 99 
density of loading, 101 
effect of thrust coefficient, 211-212 
formulas, 71-72 



impulse-weight ratio, 100 
method of obtaining high specific 
impulse, 77 

overall impulse, 100-101 
reduced impulse, 72 
specifications, 45 
thermal insulation of motors, 101 
Specifications 

antisubmarine rockets, 148-149 
chemical warfare grenade, 162 
rocket propellants, 44-45, 98, 101, 
223 

ship-to-shore rocket, 170 
Spin stabilization, theory, 215-217, 
219-222 

ballistic quantities, 215-216 
comparison with fin stabilization, 216 
effect of mallaunching, 221 
momentum, 216 
overturning moment, 219 
rocket trajectory, 221 
special purpose spinners, 221-222 
stability factor, 219-221 
yaw, 220-221 

Spin-stabilized rockets, characteristics 
accuracy, 121 

comparison with fin-stabilized, 121- 
123 

fuzes, 137 
handling, 122 

internal-burning grains, 47-48, 50 
launchers, 147 

military requirements, 121-123 
payload, 122 

simplicity and cheapness, 122 
tube design, 249-250 
velocity, 118-120 
versatility, 122 

Spin-stabilized rockets, exterior ballis- 
tics, 288-306 
air flight, 305 

ballistic constants, 289-290 
comparison with fin-stabilized, 216, 
289, 306 

effect of mallaunching, 294-296 
force system, 288-289, 298 
gravity effect, 288, 291-293, 297-298 
nutation, 288 

orientation curves, 289-291 
overturning moment, 289-291 
range calculations, 304-305 
terminal ballistics, 306 
wind effect, 296-297 
Spin-stabilized rockets, stability, 298- 
304 

causes of dispersion, 301 
effect of elevation angle, 298-299 
effect of jet malalignment, 303-304 
effect of wind, 301 
Magnus force, 299-301 



380 


INDEX 


mallaunching, 302-303 
optimum spin, 304-305 
theory, 219-221 
unbalance, 301-303 
underwater, 122 
yaw, 220-221, 298-299 
Spin-stabilized rockets, types 

3.5- in., 196-199 
3R1; 196 

4.5- in., 91 

5.0-in., 154-155, 172-173, 199-207 
aircraft, 122, 147, 203-204 
chemical spinner, 203 
general purpose aircraft spinner, 220 
high-capacity spinner, 119, 203, 
289-290 

high-velocity spinner, 201-202 
pyrotechnic, 203 
smoke spinner, 203 
SSR 

see Spin-stabilized rockets 
Stability of rockets, 219-220 
see also Yaw of rockets 
design considerations, 220 
formula, 219 

grain considerations, 231-232 
Magnus force, 299-301 
measurement, 217-218 
specifications for propellants, 44, 101 
theory, 217 
underwater, 122 
Stabilizers for rocket propellants 
centralite, 83, 102 
diphenylamine, 61, 69 
ethyl centralite, 44-45, 61, 69 
Steel rocket launcher, 185 
Stellite rocket nozzles, 258 
Step accelerometer, 28-32 
Subcaliber aircraft rocket, 176-179 
fins, 179 
heads, 179 
igniter design, 53 
launchers, 179 
lugs, 179 

nozzle, 176, 178, 252 
propellant grain, 176-177 
purpose, 176 

types and designations, 177 
Subcaliber rockets for surface warfare, 
163 

Superbazooka (rocket), 91 
Surface warfare rockets 

see Fin-stabilized rockets for surface 
warfare 

T-2 double-base rocket propellant, 
43-44, 80-81 

T-32 rocket launcher, 158 
T-40 rocket launcher, 151 
T-59 rocket, 91 


Target rockets for antiaircraft training 
advantages, 158-161 
designations and types, 161 
electrical contacts, 161 
fins, 160-161 
launchers, 161 
motor, 159-160 

Temperature in rocket propellants, 
223-225, 234-237, 259-261 
dependence on motor wall thickness, 
246 

effect on arming rocket fuze, 130 
effect on burning rate, 42-44, 85 
effect on gas velocity, 223-224 
effect on motors, 235-237, 244-246 
effect on nozzle erosion, 259 
effect on performance, 214 
effect on pressure, 225, 228 
limits, 46-47 

low temperature, 234-235 
requirements, 101 
use of blowout disks, 260-261 
variation of tensile -strength, 246 
Thermodynamics of rocket propellants, 
71-77 

attainability of high specific impulse 
fuels, 77 

calculation of gas properties, 74-75 
calculation of specific impulse, 71-72 
deviations of static measurements 
from theoretical values, 76 
discharge coefficient, 72-74, 96, 226 
effect of roughness, 77 
effective gas velocity, 71-72 
formula, 74 
fuel properties, 71-72 
heat loss, 76 
incomplete reaction, 76 
powder loss, 76-77 
thrust coefficient, 74 
3.25-in. aircraft rocket 

composition of steel in motor, 247 
nozzle, 252 

tube bending in motor, 249 

3.5-in. fin-stabilized rockets, 170-176 
center of mass, 217-218 
development history, 170-171 
fuzes, 175 

head shapes, 128, 175 
launchers and service use, 175 
propellant grain, 170 
skirts on head, 126 
tests with ballistite, 170 
types, 176 

3.5-in. fin-stabilized rockets, motor de- 
sign, 171-175 
3A9 motor, 171 
caps, 174 

electrical contacts, 174 
grids, 173 


lug bands, 173 

Mark 6 (3A12) motor, 172 

Mark 7 (3A16) motor, 172-173 

motor threads, 248 

nozzle design, 172-173 

tails, 174 

3.5-in. spin-stabilized rockets, 1969-19 
fuzes, 198 
grain, 196-197 
grid, 197 

head and motor tubes, 197-198 
igniter, 197 
launchers, 199 

nozzle plate and ring, 197-198 
seals, 198 
types, 198 

3A9 rocket motor, 171 
3A12 rocket motor, 172 
3A16 rocket motor, 172-173 
3R1 rocket, 196 

Thrust coefficient of rocket propellants 
effect of momentum and impulse, 
211-212 

effect of pressure, 213-214 
formula, 213-214 

Tin plate rocket igniters, 192-193, 
241-242 
Tiny Tim 

see 11 .75-in. aircraft rocket 
Torpedoes, 13-36 
Mark 13; 13-15, 18-19 
Mark 25; 15 
shroud ringtail, 14-15 
Torpedoes, launching tests, 21-36 
accelerometers, 28-32 
acoustic range, 22-24 
attitude (definition), 35-36 
damage instruments, 32 
deceleration, measuring equipment, 
25-27 

deviation (definition), 35-36 
drag ring, 34 
dummy torpedoes, 22-24 
entry angle, 21 
hydro pressure plugs, 33-35 
launching equipment, 21-24 
orientation recorders, 35-36 
pitch (definition), 35 
underwater photography, 27-28 
velocity-time curves, 26 
yaw (definition), 35-36 
Torpedoes, water entry, 16-20 
cavity stage, 17 

correlation between model and pro- 
totype, 20 

drag coefficient, 16-17 
effect of head shape, 18 
flow stage, 16 
immersion stage, 17 
moment of inertia, 19 


INDEX 


381 


sitch angle, 18 
phock stage, 16-17 
transition stage, 17 
trim studies, 19 

Trajectory of rockets, 272-276, 282-287 
air flight, 214-215, 274-276, 282, 305 
deviation, 267 

effect of air drag, 214-215, 270-271, 
288, 298 

effect of launcher, 142-144, 275-276 
. ground-fired, 273-274 
theory, 221 
underground, 285-287 
underwater, 282-285 
Tree-type rocket launcher, 143 
Triacetin for rocket propellants, 69, 83, 
102 

T-slot rocket launcher, 141, 175 
Tube rocket launchers, 138 
Tungsten rocket nozzles, 257-258 
2.25-in. aircraft rocket 

see Subcaliber aircraft rocket 

Underground trajectory of fin-stabil- 
ized rockets, 285-287 
Underwater ballistic studies, 8-12 
factors affecting model behavior, 
11-12 

scaled models, 8-10 
water entry of projectiles, 10 
Underwater missiles 

antisubmarine bombs, 3-8, 140-141, 
148 


torpedoes, 13-36 

Underwater photography of torpedoes, 
27-28 

Underwater trajectory of fin-stabilized 
rockets, 282-285 
cross force, 282-283 
method of controlling, 127-128 
stability, 122 

tactical effectiveness, 284-285 

University of Minnesota 

burning strand method of studying 
rocket propellants, 82-83, 101 
thermodynamics of rockets, 65, 71- 
77 

University of Wisconsin 

burning strand method of studying 
rocket propellants, 82-83, 101 
preparation of rocket propellant 
grains, 94 

thermodynamics of rocket propel- 
lants, 71-77 

VAR (vertical antisubmarine rockets) 
see Retro rockets 

Velocity of rockets, 118-121 
angular, 275-276 

comparison with machine gun bullet, 
274 

effect of motor weight, 120 r 
fin-stabilized. 118-119 
formula, 71 

spin-stabilized, 118-120 

Velocity-time curves, torpedo, 26 


Vented vessel for studying rocket pro- 
pellants, 81-82 

Vertical antisubmarine rockets 
see Retro rockets 

VFB / (vertical flare bombs), 169-170 
VFR '(vertical flare rockets), 169-170 
Vicar (rocket), 93 

Water Entry and Underwater Ballistics 
of Projectiles (report), 8-12 
Water entry of torpedoes 
see Torpedoes, water entry 
White Whizzer (rocket motor), 185 
Wind, effect on rocket trajectory 
during burning, 221 
fin-stabilized, 272-274 
spin-stabilized, 296-297, 301 
Window rockets (antiradar) 
base fuzes, 136-137 
motor, 168 

Yaw of rockets 
angle, 267 

effect on air trajectory, 275 
fin-stabilized, 218-219 
formula, 221 

spin-stabilized, 220-221, 298-299 
Yaw of torpedo, definition, 35-36 

Zero-length rocket launchers, 142-144, 
195 

effect on trajectory, 275-276 
Mark 5; 142 
tree-type, 143 


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